diff --git "a/logs/app.log" "b/logs/app.log" --- "a/logs/app.log" +++ "b/logs/app.log" @@ -1610,3 +1610,56462 @@ At the time of the incident, about 38 pounds of water was available in the potab ------ 2025-04-03 at 19:35:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:35:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 ascent stage helium pressure rise +2025-04-03 at 19:35:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:35:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: challenger spacecraft electrical anomalies investigation +2025-04-03 at 19:35:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-03 at 19:35:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: oxygen tank usage Apollo +2025-04-03 at 19:35:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 19:35:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:35:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 cryogenic pressure anomaly +2025-04-03 at 19:35:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 19:35:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: challenger spacecraft transition mode electrical failure +2025-04-03 at 19:35:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ + +2025-04-03 at 19:35:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 tank failure investigation +2025-04-03 at 19:35:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:35:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:35:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank cooling system +2025-04-03 at 19:35:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 19:35:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transition mode electrical failure transformer failure +2025-04-03 at 19:35:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-03 at 19:35:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 mission issues tank +2025-04-03 at 19:35:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 19:35:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:35:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank conditioning protocol +2025-04-03 at 19:35:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 19:35:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: induction potentiometer failure analysis potential implications +2025-04-03 at 19:35:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 19:35:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:35:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank safety testing Apollo +2025-04-03 at 19:35:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 19:35:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: three coil system C-axis induction potentiometer operation +2025-04-03 at 19:35:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 19:35:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:35:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: antenna operator C-axis and A-axis voltage interactions +2025-04-03 at 19:35:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-03 at 19:35:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:35:21 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:35:21 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:35:21 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, True, True, False, False] +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_correctness:82 - Student lengths: [531, 462, 531, 807, 558, 326] +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [27, 27, 27, 27, 27, 27] +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_correctness:84 - Average student length: 535.83 +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 27.00 +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_correctness:86 - Length ratio: 19.85 +2025-04-03 at 19:35:21 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:35:21 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.562 Âą 0.325 +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 6.50 Âą 6.73 +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 20, 9, 6, 2, 2] +2025-04-03 at 19:35:21 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 19:35:21 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +a. Some of the radioluminescent disks were broken. b. The apex cover was broken on the extravehicular handle side. c. The docking ring was burned and broken. d. The right--hand roll thruster was blistered. e. A yellowish/tan film existed on the outside of the hatch window, left and right rendezvous windows, and the right-hand window. f. The interior surfaces of the command module were very damp and cold, assumed to be condensation; there was no pooling of water on the floor. . Water samples could not be taken from the spacecraft tanks (discussed in section 5.8). h. The postlanding ventilation exhaust valve was open and the inlet valve was closed; the postlanding ventilation valve unlock handle was apparently jammed between the lock and unlock positions (section 14.l.2). i. There was more and deeper heat streaking in the area of the compression and shear pads than has been normally observed. + +11.0 EXPERIMENTS + +11.1 ATMOSPHERIC ELECTRICAL PHENOMENA +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Figure 14-6.- Recorded signal strengths during high-gain antenna operation. + +The only component identified with a failure mode that would produce a shift in the scan-limit functions and a shift change is the C-axis induction potentiometer located in the antenna. This potentiometer is used to provide a voltage proportional to the C-axis angular orientation and consists of three separate coils, each with symmetrical winding on opposite sides of the rotor or stator. These coils include the primary winding on the stator, the compensation or bias winding on the stator, and the linear output winding located on the rotor. The bias winding is used to shift the normal +70 degrees linear output to a new linear output over the range of from minus l0 to plus l30 degrees. + +The voltages for the C-axis induction potentiometer and the A-axis function generator, also located in the antenna, add together in the +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Hy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Consumed Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the incident Tank 1 21.9 255.0 Tank 2 22.3 242.0 Totals 44.2 497.0 + +7.1.4 Oxy gen + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization.bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately l4 poumds of water from the command module to the lunar module for drinking and food preparation. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 19:35:21 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...', 'Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...', 'Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...', 'Result 1:\nHy drogen, lb Oxygen, 1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 ...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\na. The mission was aborted because of the total loss of primary oxygen in the...'] +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:35:21 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:35:21 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:35:21 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.559, Max reward: 0.990 +2025-04-03 at 19:35:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:35:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "command module pilot apollo 7" +2025-04-03 at 19:35:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:35:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: John Young or Michael Collins +2025-04-03 at 19:35:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ + +2025-04-03 at 19:35:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:35:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo 7 command module pilot" +2025-04-03 at 19:35:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:35:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 17 astronaut crew +2025-04-03 at 19:35:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:35:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:35:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo 7 crew member" +2025-04-03 at 19:35:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 19:35:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:35:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo 7 command module pilot rubella" +2025-04-03 at 19:35:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:35:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:35:32 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:35:32 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:35:32 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_correctness:82 - Student lengths: [551, 86, 325, 289, 214, 277] +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [20, 20, 20, 20, 20, 20] +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_correctness:84 - Average student length: 290.33 +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 20.00 +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_correctness:86 - Length ratio: 14.52 +2025-04-03 at 19:35:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:35:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.217 Âą 0.318 +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.00 Âą 1.53 +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [4, 0, 0, 0, 2, 0] +2025-04-03 at 19:35:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:35:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi..."] +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +2025-04-03 at 19:35:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +2025-04-03 at 19:35:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +2025-04-03 at 19:35:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: ++21.0 +3.0 +------ +Result 2: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ + +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:35:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\n+21.0 +3.0\n------\nResult 2:\nTABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY\n\nMagazin...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +2025-04-03 at 19:35:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:35:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:35:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:35:32 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.121, Max reward: 0.725 +2025-04-03 at 19:35:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:35:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"space shuttle main engine accelerometer bias launch" +2025-04-03 at 19:35:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:35:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: accelerometer bias Apollo 11 translunar injection +2025-04-03 at 19:35:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 19:35:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:35:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: critical mission parameters Apollo 11 +2025-04-03 at 19:35:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:35:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:35:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was apo apollo 11 translunar injection +2025-04-03 at 19:35:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 19:35:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:35:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: initial aops accelerometer bias +2025-04-03 at 19:35:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:35:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:35:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: for more information on the initial accelerometer bias. + +initial accelerometer bias data apollo 11 +2025-04-03 at 19:35:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 19:35:45 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:35:45 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:35:46 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_correctness:82 - Student lengths: [205, 231, 353, 468, 1852, 201] +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [6, 6, 6, 6, 6, 6] +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_correctness:84 - Average student length: 551.67 +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 6.00 +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_correctness:86 - Length ratio: 91.94 +2025-04-03 at 19:35:46 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:35:46 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.137 Âą 0.198 +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.33 Âą 4.78 +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 0, 0, 1, 13, 0] +2025-04-03 at 19:35:46 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +2025-04-03 at 19:35:46 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +2025-04-03 at 19:35:46 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +2025-04-03 at 19:35:46 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 19:35:46 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\n1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJEC...', 'Result 1:\n8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH CO...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...'] +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +2025-04-03 at 19:35:46 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 1.0 +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.167 +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:35:46 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.033, Perfect scores: 0/6 +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:35:46 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:35:46 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.046, Max reward: 0.200 +2025-04-03 at 19:35:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:35:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "NASA mission service propulsion system propellant load" +2025-04-03 at 19:35:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:35:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:35:51 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:35:51 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:35:51 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 1/6 answers correct +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_correctness:82 - Student lengths: [230, 345, 504, 247, 282, 503] +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_correctness:84 - Average student length: 351.83 +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 5.00 +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_correctness:86 - Length ratio: 70.37 +2025-04-03 at 19:35:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_format:228 - Responses ending properly: 0/6 +2025-04-03 at 19:35:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.058 Âą 0.130 +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.17 Âą 0.37 +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 0, 1, 0, 0, 0] +2025-04-03 at 19:35:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 19:35:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 19:35:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:35:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nCommand module.- Command module reaction control system propellant usages can...'] +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 19:35:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 19:35:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 19:35:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:35:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:35:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:35:51 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.033, Max reward: 0.200 +2025-04-03 at 19:35:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:35:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "sustainable materials for future vehicle interiors" +2025-04-03 at 19:35:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 19:35:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Manufacturers use carbon fiber or both carbon fiber and Kevlar for shades in vehicles? +2025-04-03 at 19:35:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:35:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: automotive shade materials future +2025-04-03 at 19:35:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:35:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: future vehicle exterior shade material +2025-04-03 at 19:35:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:35:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:35:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "sustainable materials for vehicle shading, future transportation" +2025-04-03 at 19:35:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 19:35:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: elongation Aclar shades +2025-04-03 at 19:35:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:35:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:35:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ductile Aclar shade material +2025-04-03 at 19:35:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:35:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:36:00 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:36:00 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:36:00 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, True] +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_correctness:82 - Student lengths: [396, 423, 470, 318, 422, 456] +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_correctness:84 - Average student length: 414.17 +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 5.00 +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_correctness:86 - Length ratio: 82.83 +2025-04-03 at 19:36:00 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:36:00 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.308 Âą 0.241 +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.17 Âą 1.07 +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [2, 0, 1, 0, 1, 3] +2025-04-03 at 19:36:00 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +2025-04-03 at 19:36:00 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +2025-04-03 at 19:36:00 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:36:00 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.067, Perfect scores: 0/6 +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.17 +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:36:00 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:36:00 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.217, Max reward: 0.590 +2025-04-03 at 19:36:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:36:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: total oxygen supply in a surge tank incident +2025-04-03 at 19:36:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 19:36:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ott {"challenger shuttle oxygen tank loss" and total oxygen supply in tank} +2025-04-03 at 19:36:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 19:36:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was oxygen supply in the surge tank in a notable incident? +2025-04-03 at 19:36:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 19:36:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:36:07 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:36:07 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:36:07 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_correctness:82 - Student lengths: [321, 627, 575, 643, 256, 144] +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [11, 11, 11, 11, 11, 11] +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_correctness:84 - Average student length: 427.67 +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 11.00 +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_correctness:86 - Length ratio: 38.88 +2025-04-03 at 19:36:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_format:228 - Responses ending properly: 0/6 +2025-04-03 at 19:36:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.175 Âą 0.175 +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.50 Âą 0.50 +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 1, 0, 1, 0, 1] +2025-04-03 at 19:36:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +2025-04-03 at 19:36:07 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +2025-04-03 at 19:36:07 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +2025-04-03 at 19:36:07 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 19:36:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.033, Perfect scores: 0/6 +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:36:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:36:07 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.100, Max reward: 0.200 +2025-04-03 at 19:36:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:36:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"The unit of measurement for altitude above the oblate oblate spheroid Earth model" +2025-04-03 at 19:36:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 19:36:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "units of measurement for altitude above oblate spheroid earth" +2025-04-03 at 19:36:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ +Result 2: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ + +2025-04-03 at 19:36:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "altitude above oblate earth model unit" +2025-04-03 at 19:36:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ + +2025-04-03 at 19:36:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: + +2025-04-03 at 19:36:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:36:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:36:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What unit of measurement is used for altitude above the Earth in space navigation, specifically for maximum altitude? +2025-04-03 at 19:36:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ + +2025-04-03 at 19:36:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "determining unit of maximum altitude geosynchronous satellites and clouds" +2025-04-03 at 19:36:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 19:36:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:36:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "altitude unit definition cloud height geosynchronous satellites" +2025-04-03 at 19:36:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 19:36:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:36:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "cloud height unit definition geosynchronous satellite altitude" +2025-04-03 at 19:36:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 19:36:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:36:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "altitude unit definition cloud height reference Apollo 6 reference 5" +2025-04-03 at 19:36:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ +Result 2: +Marshall Space Flight Center, Kennedy Space Center, Manned Spacecraft Center: Analysis of Apollo l2 Lightning Incident, MSC-01540. February 1970. + +ICSU/IUGG Committee on Atmospheric Sciences: Report of the Study Conference on the Global Atmospheric Research Program, 1967. + +Bulletin of the American Meteorological Society, Vol. 50, No. 7: Cloud Height Contouring from Apollo 6 Photography, by V. S. Whitehead, I. D. Browne, and J. G. Garcia. 1969. + +Defense Supply Agency, Washington, D. C.: Military Standardization Handbook_ Optical Design, MIL HDBK-14l. 1962. + +NASA Headquarters: Apollo Flight Mission Assignments. OMSF M-D MA500-11 (SE 010-000-1). 0ctober 1969. + +Manned Spacecraft Center: Mission Requirement, H-2 Type Mission (Lunar Landing). SPD9-R-053. November 10, 1969. + +APOLLO SPACECRAFT FLIGHT HISTORY + +(Continued from inside front cover) +------ + +2025-04-03 at 19:36:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:36:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "altitude unit definition in NASA Apollo mission reports" +2025-04-03 at 19:36:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:36:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:36:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "altitude unit definition Apollo 11" +2025-04-03 at 19:36:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:36:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:36:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo mission altitude unit definition" +2025-04-03 at 19:36:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ¡¡īŋŊīŋŊ¡ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:36:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:36:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "altitude unit definition Apollo 11 technical report" +2025-04-03 at 19:36:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:36:27 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:36:27 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:36:27 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, False, False] +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_correctness:82 - Student lengths: [626, 2049, 348, 967, 721, 643] +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_correctness:84 - Average student length: 892.33 +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 4.00 +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_correctness:86 - Length ratio: 223.08 +2025-04-03 at 19:36:27 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:36:27 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.358 Âą 0.323 +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.17 Âą 3.13 +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [2, 9, 1, 0, 0, 1] +2025-04-03 at 19:36:27 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ + +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ +Result 2: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ + +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ +Result 2: +Marshall Space Flight Center, Kennedy Space Center, Manned Spacecraft Center: Analysis of Apollo l2 Lightning Incident, MSC-01540. February 1970. + +ICSU/IUGG Committee on Atmospheric Sciences: Report of the Study Conference on the Global Atmospheric Research Program, 1967. + +Bulletin of the American Meteorological Society, Vol. 50, No. 7: Cloud Height Contouring from Apollo 6 Photography, by V. S. Whitehead, I. D. Browne, and J. G. Garcia. 1969. + +Defense Supply Agency, Washington, D. C.: Military Standardization Handbook_ Optical Design, MIL HDBK-14l. 1962. + +NASA Headquarters: Apollo Flight Mission Assignments. OMSF M-D MA500-11 (SE 010-000-1). 0ctober 1969. + +Manned Spacecraft Center: Mission Requirement, H-2 Type Mission (Lunar Landing). SPD9-R-053. November 10, 1969. + +APOLLO SPACECRAFT FLIGHT HISTORY + +(Continued from inside front cover) +------ + +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:36:27 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe determination of the wind field in the atmosphere is one of the prime req...', 'Result 1:\ngathered from pairs of geosynchronous satellites located l0 to 20 degrees apa...', 'Result 1:\ngathered from pairs of geosynchronous satellites located l0 to 20 degrees apa...', 'Result 1:\ngathered from pairs of geosynchronous satellites located l0 to 20 degrees apa...', 'Result 1:\nThe determination of the wind field in the atmosphere is one of the prime req...', 'Result 1:\nBecause an inflight anomaly in the cryogenic oxygen supply required an abort ...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...'] +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ + +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 19:36:27 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 19:36:27 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:36:27 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\n(section ll.3).\n------\nResult 2:\n+21.0 +3.0\n------\n...'] +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:36:27 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:36:27 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:36:27 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.212, Max reward: 0.875 +2025-04-03 at 19:36:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:36:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: duration of nominal S-band communications NASA Mars Curiosity Rover +2025-04-03 at 19:36:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-03 at 19:36:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how long did nominal S-band communications last after lunar module actuation +2025-04-03 at 19:36:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 19:36:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:36:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: duration of nominal S-band communications until nominal S-band power-off +2025-04-03 at 19:36:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-03 at 19:36:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: duration of nominal s-band communications after lunar module undocking +2025-04-03 at 19:36:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 19:36:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:36:35 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:36:35 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:36:35 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_correctness:82 - Student lengths: [70, 115, 160, 643, 766, 135] +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [22, 22, 22, 22, 22, 22] +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_correctness:84 - Average student length: 314.83 +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 22.00 +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_correctness:86 - Length ratio: 14.31 +2025-04-03 at 19:36:35 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_format:228 - Responses ending properly: 0/6 +2025-04-03 at 19:36:35 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.167 Âą 0.236 +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.67 Âą 0.94 +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 0, 2, 0, 0, 2] +2025-04-03 at 19:36:35 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +2025-04-03 at 19:36:35 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +2025-04-03 at 19:36:35 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +2025-04-03 at 19:36:35 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +2025-04-03 at 19:36:35 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:36:35 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.17 +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:36:35 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:36:35 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.218, Max reward: 0.710 +2025-04-03 at 19:36:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:36:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: type "insulation in S-11 S-II space stage +2025-04-03 at 19:36:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 19:36:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What type of insulation was used in the S-II stage of a Space Shuttle +2025-04-03 at 19:36:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 19:36:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "S-1l stage insulation material +2025-04-03 at 19:36:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 19:36:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is S-1 stage? +2025-04-03 at 19:36:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 19:36:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Spacecraft S1I stage insulation type" +2025-04-03 at 19:36:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:36:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:36:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: thermal insulation S-II rocket stage +2025-04-03 at 19:36:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 19:36:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: thermal protection system of S-IB saturn ib +2025-04-03 at 19:36:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 19:36:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:36:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: insulation material S-II rocket stage +2025-04-03 at 19:36:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 19:36:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:36:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: design of spray foam insulation sponge +2025-04-03 at 19:36:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 19:36:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:36:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: degradation of vacuum insulation +2025-04-03 at 19:36:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 19:36:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:36:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: defective insulation in S-II rocket stage +2025-04-03 at 19:36:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 19:36:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:36:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apex cover jettison system S-II +2025-04-03 at 19:36:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:36:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:36:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: gas leak in S-II rocket stage apex cover jettison system +2025-04-03 at 19:36:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:36:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:36:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Teflon ring in S-II rocket stage gas leak +2025-04-03 at 19:36:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 19:36:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:36:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Teflon ring O-ring +2025-04-03 at 19:36:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 19:36:49 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:36:49 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:36:49 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1210, 194, 313, 248, 249, 61] +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_correctness:84 - Average student length: 379.17 +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_correctness:86 - Length ratio: 37.92 +2025-04-03 at 19:36:49 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:36:49 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.417 Âą 0.282 +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.50 Âą 3.40 +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [10, 0, 1, 1, 2, 1] +2025-04-03 at 19:36:49 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +2025-04-03 at 19:36:49 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 19:36:49 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...'] +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:36:49 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.033, Perfect scores: 0/6 +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:36:49 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:36:49 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.384, Max reward: 0.863 +2025-04-03 at 19:36:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:36:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: product of inertia of Lunar module before cryogenic oxygen tank ignition +2025-04-03 at 19:36:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 19:36:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module product of inertia Apollo 11 weight acrylic glass cyrogenic propellant tank fuel cryogenic tank +2025-04-03 at 19:36:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 19:36:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module inertial dynamics by NASA" +2025-04-03 at 19:36:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 19:36:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: product of inertia lunar module cryogenic tank +2025-04-03 at 19:36:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 19:36:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module acceleration theory product of inertia +2025-04-03 at 19:36:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 19:36:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:37:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module inertia before cryogenic oxygen tank ignition +2025-04-03 at 19:37:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 19:37:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moment arm distance for translational inertia +2025-04-03 at 19:37:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:37:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module spin axis product of inertia +2025-04-03 at 19:37:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:37:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module product of inertia cryogenic tank ignition +2025-04-03 at 19:37:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 19:37:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:37:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module inertia product before oxygen tank ignition 1969 +2025-04-03 at 19:37:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:37:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: `gyroaxiselectricite mapping surveyed intruction +2025-04-03 at 19:37:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 ArcturugīŧŒ36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ + +2025-04-03 at 19:37:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module inertial measurment unit heaters effects on product of inertia +2025-04-03 at 19:37:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 19:37:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:37:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module state vector error matrix coefficients 1969 +2025-04-03 at 19:37:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:37:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: gyrocompassing required alignment parameters +2025-04-03 at 19:37:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 19:37:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:37:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 lunar module dynamics analysis +2025-04-03 at 19:37:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:37:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: signal processingasy method error estimation +2025-04-03 at 19:37:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:37:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:37:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apo11 lunar module physical parameters +2025-04-03 at 19:37:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-03 at 19:37:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft control principles, offset velocity, ft/sec +2025-04-03 at 19:37:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 19:37:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:37:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module inertia values before oxygen tank ignition +2025-04-03 at 19:37:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 19:37:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lnterface between accelerometer and gyro measurements +2025-04-03 at 19:37:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:37:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:37:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module inertial measurement unit bias after heater power-up +2025-04-03 at 19:37:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:37:17 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:37:17 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:37:17 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_correctness:82 - Student lengths: [2009, 489, 1806, 1066, 480, 647] +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_correctness:84 - Average student length: 1082.83 +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 5.00 +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_correctness:86 - Length ratio: 216.57 +2025-04-03 at 19:37:17 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:37:17 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.487 Âą 0.288 +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 7.00 Âą 9.73 +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [8, 1, 28, 0, 2, 3] +2025-04-03 at 19:37:17 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:37:17 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...', 'Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...', 'Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nA.2 LUNAR MODULE\n------\nResult 2:\nSpacecraft mass properties for the Apollo l...', 'Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...', 'Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...'] +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 19:37:17 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\n1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJEC...'] +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 ArcturugīŧŒ36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ + +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:37:17 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...', 'Result 1:\nupdated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 ...', 'Result 1:\nupdated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 ...', 'Result 1:\nTable 5.6-I is a summary of gyro drift measurements deduced from inflight ali...', 'Result 1:\nAfter completing the maneuver to the attitude required for the final midcours...', 'Result 1:\nEarth-centered inertial coordinates. Bystem. **Change in velocity showm in bo...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...'] +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 19:37:17 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:37:17 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...', 'Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...'] +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 1.0 +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.167 +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:37:17 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.200, Perfect scores: 0/6 +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.50 +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:37:17 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:37:17 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.297, Max reward: 0.634 +2025-04-03 at 19:37:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:37:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"standard deviation of scale factor X-axis accelerometer structural health monitoring +2025-04-03 at 19:37:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:37:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: X-axis accelerometer microPiezoelectrics standard deviation of scale factor error +2025-04-03 at 19:37:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:37:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: high-precision accelerometer X-axis scale factor error standard deviation +2025-04-03 at 19:37:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-03 at 19:37:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the technical definition of scale factor error for an acceleration sensor or accelerometer? +2025-04-03 at 19:37:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 19:37:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How does the standard deviation of the scale factor error for X-axis accelerometers in MEMS-based devices compare across different operating temperatures? +2025-04-03 at 19:37:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:37:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:37:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: accelerationometer&scale factor error&X-axis acceleration drift &accelerometer bias&scale factor error +2025-04-03 at 19:37:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:37:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the standard deviation of acceleration error in the X-axis accelerometer? +2025-04-03 at 19:37:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:37:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:37:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: specific x-axis accelerometer standard deviation of scale factor error +2025-04-03 at 19:37:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:37:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the standard deviations for the individual scale factor error values of the X, Y, and Z axes? +2025-04-03 at 19:37:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 19:37:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:37:33 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:37:33 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:37:33 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, True] +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_correctness:82 - Student lengths: [232, 347, 612, 1012, 711, 239] +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_correctness:84 - Average student length: 525.50 +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 4.00 +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_correctness:86 - Length ratio: 131.38 +2025-04-03 at 19:37:33 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:37:33 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.392 Âą 0.221 +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.50 Âą 1.12 +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [1, 0, 3, 1, 3, 1] +2025-04-03 at 19:37:33 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +2025-04-03 at 19:37:33 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 īŋŊīŋŊ -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:37:33 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...'] +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:37:33 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:37:33 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:37:33 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.326, Max reward: 0.691 +2025-04-03 at 19:37:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:37:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spaceplane capable of lifting command module +2025-04-03 at 19:37:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +Upon completion of photography, the two docked vehicles were maneuvered back to the service module separation attitude in preparation for the command module alignment. Star observation through the command module optics in this attitude was poor because of light reflecting from the lunar module, and the Conmander varied the pitch attitude by approximately 20 degrees in an attempt to improve star visibility. These attitude excursions, however, were not effective, and the spacecraft was returned to the original separation attitude for the command module alignment. + +8.9.5 Command Module Activation +------ + +2025-04-03 at 19:37:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: crew transportation to orbit vehicles +2025-04-03 at 19:37:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:37:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space shuttle orbiters +2025-04-03 at 19:37:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:37:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Spacecraft Command Module Lifter +2025-04-03 at 19:37:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:37:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:37:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft carrying crew command module +2025-04-03 at 19:37:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 19:37:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 1 pitch maneuver +2025-04-03 at 19:37:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:37:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA spacecraft transport systems +2025-04-03 at 19:37:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:37:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V lifting capacity +2025-04-03 at 19:37:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-03 at 19:37:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:37:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: crewed spaceflight command module separation +2025-04-03 at 19:37:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-03 at 19:37:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 01 crew members +2025-04-03 at 19:37:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:37:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Lunar Module +2025-04-03 at 19:37:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:37:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V vs jet aircraft comparison +2025-04-03 at 19:37:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 19:37:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:37:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar mission crewed spacecraft command module +2025-04-03 at 19:37:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:37:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 reaction control system +2025-04-03 at 19:37:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:37:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 3 +2025-04-03 at 19:37:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:37:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:37:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module command module separation +2025-04-03 at 19:37:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:37:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 mission overview +2025-04-03 at 19:37:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:37:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module +2025-04-03 at 19:37:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:37:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:37:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 vs Apollo 11 Lunar Module reaction control system performance +2025-04-03 at 19:37:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:37:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 hydro pressure considerations +2025-04-03 at 19:37:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 19:37:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:37:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 vs Apollo 11 Lunar Module analysis comparison +2025-04-03 at 19:37:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:37:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 helium pressure issues +2025-04-03 at 19:37:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 19:37:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:37:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 vs Apollo 11 evaluation supplement differences +2025-04-03 at 19:37:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:37:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module hydrogen tank pressure +2025-04-03 at 19:37:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 19:37:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:37:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 cancelled analysis supplements +2025-04-03 at 19:37:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:37:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apodouna-8 +2025-04-03 at 19:37:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 19:37:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:37:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar landing +2025-04-03 at 19:37:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:37:59 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:37:59 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:37:59 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_correctness:82 - Student lengths: [332, 330, 2038, 1984, 216, 381] +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [18, 18, 18, 18, 18, 18] +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_correctness:84 - Average student length: 880.17 +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 18.00 +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_correctness:86 - Length ratio: 48.90 +2025-04-03 at 19:37:59 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:37:59 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.425 Âą 0.340 +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 5.00 Âą 4.58 +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 2/6 +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 5, 11, 11, 0, 3] +2025-04-03 at 19:37:59 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +2025-04-03 at 19:37:59 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +Upon completion of photography, the two docked vehicles were maneuvered back to the service module separation attitude in preparation for the command module alignment. Star observation through the command module optics in this attitude was poor because of light reflecting from the lunar module, and the Conmander varied the pitch attitude by approximately 20 degrees in an attempt to improve star visibility. These attitude excursions, however, were not effective, and the spacecraft was returned to the original separation attitude for the command module alignment. + +8.9.5 Command Module Activation +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:37:59 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nstarted to sight the service module in the docking window. The lightened spac...', 'Result 1:\nThe service module was separated $\\scriptstyle4-3/4$ hours before entry , aff...', 'Result 1:\nA major flight control function, in adaition to the monitoring of systems sta...', 'Result 1:\nThe unprecedented powered-down state of the command module required generatio...', 'Result 1:\nThe service module was separated $\\scriptstyle4-3/4$ hours before entry , aff...'] +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:37:59 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nFollowing separation and translation, a manual pitch maneuver of 1.5 deg/sec ...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...'] +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:37:59 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...', 'Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...', 'Result 1:\nA.2 LUNAR MODULE\n------\nResult 2:\nd. The effectiveness of preflight crew trai...', 'Result 1:\n1.0 SUMMARY\n\nThe Apollo l3 mission, planned as a lunar landing in the Fra Mau...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\ndiscussion of the flight and associated crew activities is presented in secti...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +2025-04-03 at 19:37:59 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 19:37:59 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ["Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nAt lift-off, measured winds, both at the surface and in the region of maximum...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...'] +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:37:59 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.200, Perfect scores: 1/6 +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.33 +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 19:37:59 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:37:59 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.425, Max reward: 0.727 +2025-04-03 at 19:38:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:38:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Daughter board failure rate in Hawaii EPA nuclear-related cleanup" +2025-04-03 at 19:38:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 19:38:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"0065 Hawaii nuclear incident propellant isolation valves" + +2025-04-03 at 19:38:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 19:38:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Hawaii propellant isolation valve system decontamination incident report" +2025-04-03 at 19:38:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 19:38:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:38:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +I was unable to find any results that match the previous search query, but I will modify it to Result 1: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 19:38:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:38:07 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:38:07 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:38:07 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, False, False, False, False] +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_correctness:82 - Student lengths: [258, 785, 452, 28, 461, 193] +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [15, 15, 15, 15, 15, 15] +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_correctness:84 - Average student length: 362.83 +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 15.00 +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_correctness:86 - Length ratio: 24.19 +2025-04-03 at 19:38:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:38:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.200 Âą 0.206 +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.67 Âą 0.75 +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 1, 1, 0, 2, 0] +2025-04-03 at 19:38:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +2025-04-03 at 19:38:07 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +2025-04-03 at 19:38:07 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +2025-04-03 at 19:38:07 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 19:38:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:38:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:38:07 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.100, Max reward: 0.200 +2025-04-03 at 19:38:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:38:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "discrepancy in suit pressure measurement" +2025-04-03 at 19:38:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 19:38:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query to get more information. +search "discrepancy in suit pressure measurement" +2025-04-03 at 19:38:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 19:38:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:38:12 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:38:12 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:38:12 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_correctness:82 - Student lengths: [525, 234, 405, 244, 335, 230] +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [28, 28, 28, 28, 28, 28] +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_correctness:84 - Average student length: 328.83 +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 28.00 +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_correctness:86 - Length ratio: 11.74 +2025-04-03 at 19:38:12 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:38:12 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.117 Âą 0.165 +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.33 Âą 0.47 +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 1, 0, 1, 0, 0] +2025-04-03 at 19:38:12 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +2025-04-03 at 19:38:12 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +2025-04-03 at 19:38:12 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +2025-04-03 at 19:38:12 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +2025-04-03 at 19:38:12 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:38:12 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:38:12 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:38:12 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.067, Max reward: 0.200 +2025-04-03 at 19:38:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:38:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module undocking from command module Apollo 11 +2025-04-03 at 19:38:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 19:38:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module undocking time after moon landing +2025-04-03 at 19:38:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:38:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo lunar module undocking timeline" +2025-04-03 at 19:38:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:38:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Apollo 11 lunar module undocking time +2025-04-03 at 19:38:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 19:38:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module undocking time to earth orbit +2025-04-03 at 19:38:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:38:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:38:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module undocking time +2025-04-03 at 19:38:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 19:38:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo lunar module undocking procedure" +2025-04-03 at 19:38:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:38:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:38:20 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:38:20 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:38:20 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_correctness:82 - Student lengths: [282, 547, 715, 278, 125, 221] +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [6, 6, 6, 6, 6, 6] +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_correctness:84 - Average student length: 361.33 +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 6.00 +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_correctness:86 - Length ratio: 60.22 +2025-04-03 at 19:38:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:38:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.342 Âą 0.167 +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.17 Âą 0.69 +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [2, 1, 0, 2, 1, 1] +2025-04-03 at 19:38:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 19:38:20 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...', 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...'] +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:38:20 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...'] +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 19:38:20 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:38:20 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', 'Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...'] +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 19:38:20 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...'] +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:38:20 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...'] +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:38:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.067, Perfect scores: 0/6 +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:38:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:38:20 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.248, Max reward: 0.575 +2025-04-03 at 19:38:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:38:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: a spacecraft Ascent stage with Airbreather propulsion. +2025-04-03 at 19:38:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 19:38:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ascent stage Tank 2 oxygen consumption +2025-04-03 at 19:38:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 19:38:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +The 'Ascent' is part of Blue Origin's New Shepard rocket system. Interestingly, New Shepard's upper stage also goes by 'Ascent'. The only stage referred to as 'Tank 2' in that specific vehicle is the first stage. I will search for the first stage oxygen consumption of the New Shepard. + + +2025-04-03 at 19:38:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-03 at 19:38:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent stage tank capacity oxygen consumption Blue Origin New Shepard +2025-04-03 at 19:38:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 19:38:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Ascent stage tank 2 oxygen consumption" +2025-04-03 at 19:38:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 19:38:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:38:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Westinghouse Airbreather propulsion system Ascent stage. +2025-04-03 at 19:38:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:38:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Another search will be done to establish what happens when a BE-4 engine, powered by oxygen, goes out during the ascent phase of New Shepard, in anticipation of tube assembly issues possibly, as part of the precious fakes for BE-4 minutes prior to reaching or passing the majority of operational altitude. This reasoning point leads me to look for information relating to the BE-4 engine failure or, at the least limiting factors in oxygen weight and leaked oxygen rates. + + +2025-04-03 at 19:38:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ + +2025-04-03 at 19:38:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Blue Origin New Shepard ascent stage tank oxygen consumption specifics +2025-04-03 at 19:38:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 19:38:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:38:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Westinghouse Airbreather propellant consumption liquid hydrogen consumption during ascent +2025-04-03 at 19:38:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 19:38:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Because the BE4 and New Shepard would work directly, the possibility of power loss caused by one system failure in one stage would be a cause for redundant measures to another stage, this examines systems with a BE4 resulting in steps related to operational accumulations instead of flow rates or zone usage through continuous redundancy system to reserve critical operational systems like on BE4 engine generation current entirely. + + + + +Also, for better to establish comparisons with knowledge gathered so far, I am going to research lock on activation to complete what is available. And because it will vary both the method of engine activation and could vary the operational capability of each system and operational. Considering fluid dynamics on propellant flow the flow dynamics logic behind immersed 'redundant' backup emergency power locks in primary control. + + +2025-04-03 at 19:38:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ + +2025-04-03 at 19:38:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:38:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system helium usage +2025-04-03 at 19:38:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:38:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:38:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ascent stage total propellant consumption +2025-04-03 at 19:38:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 19:38:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:38:42 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:38:42 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:38:42 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, True, True] +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_correctness:82 - Student lengths: [277, 656, 506, 764, 45, 259] +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [3, 3, 3, 3, 3, 3] +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_correctness:84 - Average student length: 417.83 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 3.00 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_correctness:86 - Length ratio: 139.28 +2025-04-03 at 19:38:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:38:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.425 Âą 0.281 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.17 Âą 1.77 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 5, 1, 4, 2, 1] +2025-04-03 at 19:38:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +2025-04-03 at 19:38:42 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ + +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ + +2025-04-03 at 19:38:42 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nDuring the flight, the pressure in the ascent stage oxygen tank 2 increased, ...', 'Result 1:\nBecause an inflight anomaly in the cryogenic oxygen supply required an abort ...', 'Result 1:\nThe service propulsion auxiliary propellant gaging system failed prior to lau...'] +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 19:38:42 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nDuring the flight, the pressure in the ascent stage oxygen tank 2 increased, ...', 'Result 1:\nDuring the flight, the pressure in the ascent stage oxygen tank 2 increased, ...'] +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 19:38:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.133, Perfect scores: 0/6 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.33 +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:38:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:38:42 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.430, Max reward: 0.853 +2025-04-03 at 19:38:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:38:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 14 service module unexplained velocity increase +2025-04-03 at 19:38:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ + +2025-04-03 at 19:38:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar orbit s-ivb velocity increase +2025-04-03 at 19:38:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 19:38:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the S-IVB velocity increase after lunar explosion +2025-04-03 at 19:38:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 19:38:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:38:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 command module separation reaction control system anomaly +2025-04-03 at 19:38:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 19:38:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moon s-ivb impact velocity increase 19h 17m +2025-04-03 at 19:38:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 19:38:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 S-IVB impact success +2025-04-03 at 19:38:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:38:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:38:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 unexplained velocity regain after lunar module jettison +2025-04-03 at 19:38:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:38:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 S-IVB impact velocity increase explanation +2025-04-03 at 19:38:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:38:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:38:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 trans-lunar injection and velocity anomaly +2025-04-03 at 19:38:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:38:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:38:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module velocity anomaly during lunar orbit +2025-04-03 at 19:38:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:38:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:38:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module separation velocity anomaly +2025-04-03 at 19:38:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:38:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:38:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module reaction control system separation velocity +2025-04-03 at 19:38:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:38:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:38:56 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:38:56 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:38:57 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_correctness:82 - Student lengths: [738, 387, 338, 263, 414, 384] +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_correctness:84 - Average student length: 420.67 +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 8.00 +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_correctness:86 - Length ratio: 52.58 +2025-04-03 at 19:38:57 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 19:38:57 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.350 Âą 0.374 +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.00 Âą 2.52 +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [7, 2, 0, 0, 0, 3] +2025-04-03 at 19:38:57 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ + +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +14.0 ANOMALY SUMMARY + +This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission. + +14.1 COMMAND AND SERVICE MODULES + +14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure + +At approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp "bang," coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system: +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:38:57 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe structure in the forward end of the docking tunnel was reinforced to acco...', 'Result 1:\nThe pilot describing function experiment (T-029) was a success, in that data ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe unusual spacecraft configuration required that new procedures for entry b...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...'] +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +2025-04-03 at 19:38:57 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +2025-04-03 at 19:38:57 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +2025-04-03 at 19:38:57 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:38:57 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.033, Perfect scores: 0/6 +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:38:57 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:38:57 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.337, Max reward: 0.717 +2025-04-03 at 19:38:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:39:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the initial propellant consumption of the Space Shuttle Main Engine's reaction control system? +2025-04-03 at 19:39:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 19:39:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system propellant consumption +2025-04-03 at 19:39:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:39:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the initial propellant consumption of the reaction control system in a spacecraft? +2025-04-03 at 19:39:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 19:39:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:39:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the initial propellant consumption of the reaction control system? +2025-04-03 at 19:39:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 19:39:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system initial activation time +2025-04-03 at 19:39:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 19:39:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system command module propellant consumption +2025-04-03 at 19:39:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:39:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:39:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system fuel usage initial activation +2025-04-03 at 19:39:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 19:39:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module reaction control system propellant consumption +2025-04-03 at 19:39:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:39:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:39:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module reaction control system propellant consumption reaction control +2025-04-03 at 19:39:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 19:39:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:39:08 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:39:08 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:39:08 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, True, False] +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_correctness:82 - Student lengths: [27, 475, 188, 298, 205, 401] +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_correctness:84 - Average student length: 265.67 +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_correctness:86 - Length ratio: 26.57 +2025-04-03 at 19:39:08 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:39:08 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.325 Âą 0.336 +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.50 Âą 1.61 +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [2, 3, 0, 0, 4, 0] +2025-04-03 at 19:39:08 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 19:39:08 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 19:39:08 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +2025-04-03 at 19:39:08 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 19:39:08 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.267, Perfect scores: 0/6 +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.00 +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:39:08 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:39:08 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.206, Max reward: 0.463 +2025-04-03 at 19:39:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:39:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: packaging for nasal, ear, or throat drops for medical kits in aviation +2025-04-03 at 19:39:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 19:39:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what type of nasal drops will be packaged with first aid kits for flights +2025-04-03 at 19:39:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 19:39:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:39:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasal drops packaging in Apollo medical kits +2025-04-03 at 19:39:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 19:39:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasal spray nasal drops for space travel +2025-04-03 at 19:39:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 19:39:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:39:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasal drops for space travelers broken up into similar packaging as eye drops used for nasal congestion or sinus issues +2025-04-03 at 19:39:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 19:39:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:39:17 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:39:17 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:39:17 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, False, True] +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_correctness:82 - Student lengths: [336, 118, 480, 94, 493, 353] +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_correctness:84 - Average student length: 312.33 +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_correctness:86 - Length ratio: 31.23 +2025-04-03 at 19:39:17 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 19:39:17 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.192 Âą 0.274 +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.83 Âą 1.21 +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [2, 0, 0, 0, 0, 3] +2025-04-03 at 19:39:17 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 19:39:17 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 19:39:17 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 19:39:17 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 19:39:17 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +This anomaly is closed. + +14.3.3 Improper Nasal Spray Operation + +When attempts were made to use the two nasal spray bottles in the command module medical kit, no medication could be obtained from one bottle and only two or three sprays could be obtained from the other. On previous flights, there had been a tendency for the spray to be released too fast, therefore a piece of cotton was inserted in the 9-cc bottle to hold the 3 cc of medication. Chamber tests and ambient shelflife tests have indicated that this change was satisfactory. Those tests have also shown that, for best results, the bottle should be squeezed where the cotton is located. Postflight examination of the one returned bottle demonstrated satisfactory operation under normal gravity. The returned bottle still contained 2.5 cc of medication after five or six test sprays. +------ + +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:39:17 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.133, Perfect scores: 0/6 +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.17 +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 19:39:17 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:39:17 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.212, Max reward: 0.660 +2025-04-03 at 19:39:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:39:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the location of the first landing site of humans in space? +2025-04-03 at 19:39:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:39:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the location of the crash site of the first lunar landing +2025-04-03 at 19:39:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:39:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the landing site of a plane crash that received widespread media attention +2025-04-03 at 19:39:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +------ + +2025-04-03 at 19:39:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: landing site location Apollo 11 +2025-04-03 at 19:39:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 19:39:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:39:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the first manned landing site in space? +2025-04-03 at 19:39:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:39:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 landing site coordinates +2025-04-03 at 19:39:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:39:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: recoveries from Iwo Jima +2025-04-03 at 19:39:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 19:39:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 25.5°S 165.42°W +2025-04-03 at 19:39:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 19:39:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:39:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of first lunar landing site +2025-04-03 at 19:39:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 19:39:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 landing coordinates +2025-04-03 at 19:39:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 19:39:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: recovery operations astronaut command module aftermath mission +2025-04-03 at 19:39:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 19:39:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ap01 lunar landing temp +2025-04-03 at 19:39:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 19:39:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:39:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: aborted space mission recovery operations Iwo Jima +2025-04-03 at 19:39:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 19:39:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 landing site lunar module +2025-04-03 at 19:39:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:39:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:39:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: direct Apollo 13 landing site +2025-04-03 at 19:39:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:39:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:39:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Samoa and New Zealand +2025-04-03 at 19:39:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 19:39:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:39:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: USS Iwo Jima arrival +2025-04-03 at 19:39:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +After confirrning the integrity of the command module and the status of the crew, the Recovery helicopter crew attempted to recover the main parachutes with grappling hooks and flotation gear prior to their sinking. Swim l and Swim 2 helicopters arrived on scene and immediately proceeded with retrieval. Swim 2 deployed swimmers to provide flotation to the spacecraft, and Swim l deployed swimmers to retrieve the apex cover, which was located upwind of the spacecraft. The flight crew was onboard the recovery helicopter 7 minutes after they had egressed the command module, and they arrived aboard Iwo Jima at 1853 G.m.t. + +Command module retrieval took place at 2l degrees 39.l minutes south latitude and 165 degrees 20.9 minutes west longitude at 1936 G.m.t. One main parachute and the apex cover were retrieved by small boat and brought aboard. +------ + +2025-04-03 at 19:39:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:39:32 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:39:32 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:39:32 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, True, False] +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_correctness:82 - Student lengths: [136, 320, 372, 243, 509, 236] +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [17, 17, 17, 17, 17, 17] +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_correctness:84 - Average student length: 302.67 +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 17.00 +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_correctness:86 - Length ratio: 17.80 +2025-04-03 at 19:39:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:39:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.429 Âą 0.318 +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 3.67 Âą 4.03 +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [3, 3, 0, 4, 12, 0] +2025-04-03 at 19:39:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 19:39:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...', 'Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...'] +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 19:39:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nAttempt to impact the expended S-IVB stage on the lunar surface within 350 km...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...'] +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +2025-04-03 at 19:39:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +TABLE 1O.3-I.- RECOVERY SUPPORT + +Landing area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, ’Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii + +&arotal ship support $=5$ Total aircraft support $\approx$ 23 + +10.3.l Command Module Location and Retrieval +------ + +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 19:39:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe ship-based aircraft were deployed relative to the Iwo Jima and were on st...', 'Result 1:\nThe flight crew remained aboard the Iwo Jima overnight and were flown to Pago...', 'Result 1:\nThe Mission Control Center and the Manned Space Flight Network provided excel...', 'Result 1:\nThe ship-based aircraft were deployed relative to the Iwo Jima and were on st...'] +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +After confirrning the integrity of the command module and the status of the crew, the Recovery helicopter crew attempted to recover the main parachutes with grappling hooks and flotation gear prior to their sinking. Swim l and Swim 2 helicopters arrived on scene and immediately proceeded with retrieval. Swim 2 deployed swimmers to provide flotation to the spacecraft, and Swim l deployed swimmers to retrieve the apex cover, which was located upwind of the spacecraft. The flight crew was onboard the recovery helicopter 7 minutes after they had egressed the command module, and they arrived aboard Iwo Jima at 1853 G.m.t. + +Command module retrieval took place at 2l degrees 39.l minutes south latitude and 165 degrees 20.9 minutes west longitude at 1936 G.m.t. One main parachute and the apex cover were retrieved by small boat and brought aboard. +------ + +2025-04-03 at 19:39:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe command module was powered up with the three entry batteries, which had b...', 'Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...', "Result 1:\nThe Iwo Jima's position was established accurately using a satellite navigati..."] +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +2025-04-03 at 19:39:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:39:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.200, Perfect scores: 0/6 +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.50 +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:39:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:39:32 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.384, Max reward: 0.666 +2025-04-03 at 19:39:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:39:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 12 signal intensity peak vs ascent stage impact" +2025-04-03 at 19:39:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:39:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar impact intensity Apollo 12 +2025-04-03 at 19:39:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-03 at 19:39:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Celestially trajectory and lunar modules ascent intensity +2025-04-03 at 19:39:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-03 at 19:39:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 ascent stage impact signal intensity +2025-04-03 at 19:39:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:39:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 ascent stage impact signal intensity peak +2025-04-03 at 19:39:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:39:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:39:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 12 S-IVB vs ascent stage seismic signal magnitude" +2025-04-03 at 19:39:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:39:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-IVB impact peak signal intensity +2025-04-03 at 19:39:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:39:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 ascent stage impact peak signal intensity compared to S-IVB impact +2025-04-03 at 19:39:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:39:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 ascent stage impact signal intensity comparison +2025-04-03 at 19:39:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:39:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:39:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-IVB impact 10 minutes peak signal intensity equivalent to ascent stage impact +2025-04-03 at 19:39:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:39:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 ascent stage impact signal intensity equivalent in tons TNT +2025-04-03 at 19:39:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-03 at 19:39:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:39:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: TNT equivalent Apollo 12 ascent stage impact +2025-04-03 at 19:39:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:39:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:39:44 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:39:44 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:39:44 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, True, False, False, True, True] +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.67 +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_correctness:82 - Student lengths: [335, 547, 170, 514, 415, 550] +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [14, 14, 14, 14, 14, 14] +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_correctness:84 - Average student length: 421.83 +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 14.00 +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_correctness:86 - Length ratio: 30.13 +2025-04-03 at 19:39:44 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 19:39:44 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.467 Âą 0.251 +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.00 Âą 1.29 +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [2, 0, 3, 1, 4, 2] +2025-04-03 at 19:39:44 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +2025-04-03 at 19:39:44 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was l9:l3:00 G.m.t., April ll, l970. All references to mileage distance are in nautical. miles. + +The Apollo l3 mission was planned as a precision lunar landing in the Fra Mauro highlands. The most significant changes to the planned mission profile from Apollo l2 were the maneuver to impact the depleted S-IVB stage on the lunar surface and the performance of descent orbit insertion using the service propulsion system. The S-IvB impact was intended to provide seismological data sensed by the instrument left on the moon during Apollo l2. Performance of the descent orbit insertion using the service propulsion system provides a greater propellant margin in the lunar module descent propulsion system, and this reserve would have been available during the critical precision landing phase. +------ + +2025-04-03 at 19:39:44 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe analysis of the trajectory from lift-off to spacecraft/S-IVB separation w...'] +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ + +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +In prior lunar missions, the third stage has been separated from the spacecraft with the intention of entering a solar orbit through a nearmiss, or "slingshot," approach to the moon. For Apollo l3, an opportunity was available to gain further data on large-mass impact phenomena which could be derived using the seismic equipment deployed during Apollo l2. The impact of the lunar module ascent stage during Apollo 12 pointed up certain unexplained seismological events which the S-IVB impact was expected to reproduce. + +The S-IVB impacted the lunar surface at 8:09:41 p.m. e.s.t., April 14, 1970, trave1ling at a speed of 5600 miles/hr. Stage weight at the time of impact was 30 700 pounds. The collision occurred at a latitude of 2.4 degrees south and a longitude of 27.9 degrees west, which is approximately 74 miles west-northwest from the experiment station installed during Apollo l2._ The energy release from the impact was equivalent to an explosion of 7.7 tons of trinitrotoluene (TNT). +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:39:44 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.167, Perfect scores: 0/6 +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.33 +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:39:44 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:39:44 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.362, Max reward: 0.703 +2025-04-03 at 19:39:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:39:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AK 103 ammunition specifications +2025-04-03 at 19:39:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ + +2025-04-03 at 19:39:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the format of pablo picasso guernica's frame and what is it and how high is the focal point of the painting +2025-04-03 at 19:39:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Positions are extrapolated. + +The ll photographs have been normalized so that the earth is the same size in all frames. Frames 8590 and 8591 have been further enlarged. By viewing these two frames under a stereoscope, pronounced apparent relief is seen in the cloud patterns. The relief is so pronounced, in fact, that it cannot be attributed solely to height differences of clouds. It appears to result, in part, from the relative horizontal motion in the cloud fields; that is, clouds moving in the same direction as the spacecraft appear farther away than those moving in the direction opposite that of the spacecraftīŧŒ + +11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT +------ +Result 2: +according to ground instructions. +------ + +2025-04-03 at 19:39:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: blizing star stratos alignment minecraft +2025-04-03 at 19:39:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 19:39:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the meaning of 13-60-8594? +2025-04-03 at 19:39:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:39:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 13-60-8594 pattern reference " " +2025-04-03 at 19:39:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +14.2.2。 +------ + +2025-04-03 at 19:39:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the meaning of 13-60-8594 and what context it belongs to +2025-04-03 at 19:39:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +14.2.2。 +------ + +2025-04-03 at 19:39:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:39:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: backwards calculation 1966 Super Circular entry supersonic heat load +2025-04-03 at 19:39:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +At about l09 hours when the helium bottle pressure had reached approximately l937 psi, the burst diaphragm ruptured and relieved the supercritical system through a special non-propulsive vent. The predicted rupture range for this vehicle was 190o Âą 20 psia. During venting, unexpected motion was inparted to the spacecraft which disrupted the motion established for the passive thermal control mode. The vent tube for the supercritical helium tank is ported on two sides by diametrically opposed Oval-shaped holes. It was originally believed that the escaping gas would exit these holes at 9o degrees to the tube axis such that no net thrust is produced. However, the pressure distribution in the tube is such that the two gas plumes have an included angle less than 180 degrees and probably closer to 9o degrees. Therefore, the component of the gas flow along the axis of the vent tube produces a net thrust in the opposite direction which tends to induce a slight roll rate to the vehicle. Since +------ + +2025-04-03 at 19:39:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 mission altitude and alignment 13-60-8594 +2025-04-03 at 19:39:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:39:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the significance of the numbers in 13-60-8594? +2025-04-03 at 19:39:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:39:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what does 13-60-8594 represent in aviation or astronomy +2025-04-03 at 19:39:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:39:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:39:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: SpaceX Frost bite failure +2025-04-03 at 19:39:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +a. Some of the radioluminescent disks were broken. b. The apex cover was broken on the extravehicular handle side. c. The docking ring was burned and broken. d. The right--hand roll thruster was blistered. e. A yellowish/tan film existed on the outside of the hatch window, left and right rendezvous windows, and the right-hand window. f. The interior surfaces of the command module were very damp and cold, assumed to be condensation; there was no pooling of water on the floor. . Water samples could not be taken from the spacecraft tanks (discussed in section 5.8). h. The postlanding ventilation exhaust valve was open and the inlet valve was closed; the postlanding ventilation valve unlock handle was apparently jammed between the lock and unlock positions (section 14.l.2). i. There was more and deeper heat streaking in the area of the compression and shear pads than has been normally observed. + +11.0 EXPERIMENTS + +11.1 ATMOSPHERIC ELECTRICAL PHENOMENA +------ + +2025-04-03 at 19:39:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 transearth injection navigation +2025-04-03 at 19:39:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-03 at 19:39:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the GPS coordinates associated with 13-60-8594? +2025-04-03 at 19:39:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 19:39:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what altitude did the Apollo 13 spacecraft fly at +2025-04-03 at 19:39:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:39:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:39:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 altitude during splate recovery +2025-04-03 at 19:39:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:39:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: tranesarth injection maneuver object south pacific +2025-04-03 at 19:39:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 19:39:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what altitude did the Apollo 13 spacecraft reach +2025-04-03 at 19:39:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:39:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:40:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 13-60-8594 object apollo mission +2025-04-03 at 19:40:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 19:40:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the altitude at which the lunar module of the Apollo 13 mission reached +2025-04-03 at 19:40:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 19:40:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:40:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 13-60-8594 recovered object +2025-04-03 at 19:40:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ +Result 2: +The first reported electronic contact by the recovery forces was through S-band contact by Samoa Rescue 4. A visual sighting report by the Recovery helicopter was received and was followed shortly thereafter by aquisition of the recovery beacon signal by the Recovery, Photo, and Swim l helicopters. Fuel dump was noted and voice contact was made with the descending spacecraft, although no latitude and longitude data were received. The command module landed at 1807 G.m.t. and remained in the stable l flotation attitude. The flashing light was operating and the infiation of the uprighting system commenced about l0 minutes subsequent to landing. +------ + +2025-04-03 at 19:40:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:40:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 13-60-8594 mission details post recovery +2025-04-03 at 19:40:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The first reported electronic contact by the recovery forces was through S-band contact by Samoa Rescue 4. A visual sighting report by the Recovery helicopter was received and was followed shortly thereafter by aquisition of the recovery beacon signal by the Recovery, Photo, and Swim l helicopters. Fuel dump was noted and voice contact was made with the descending spacecraft, although no latitude and longitude data were received. The command module landed at 1807 G.m.t. and remained in the stable l flotation attitude. The flashing light was operating and the infiation of the uprighting system commenced about l0 minutes subsequent to landing. +------ + +2025-04-03 at 19:40:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:40:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 13-60-8594 specific contents +2025-04-03 at 19:40:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:40:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:40:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 MSC-02545 +2025-04-03 at 19:40:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 19:40:08 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:40:08 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:40:08 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, True, True] +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_correctness:82 - Student lengths: [626, 388, 1653, 324, 413, 552] +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [6, 6, 6, 6, 6, 6] +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_correctness:84 - Average student length: 659.33 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 6.00 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_correctness:86 - Length ratio: 109.89 +2025-04-03 at 19:40:08 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 19:40:08 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.642 Âą 0.235 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 4.67 Âą 2.43 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 2/6 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [5, 1, 9, 3, 5, 5] +2025-04-03 at 19:40:08 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ + +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +At about l09 hours when the helium bottle pressure had reached approximately l937 psi, the burst diaphragm ruptured and relieved the supercritical system through a special non-propulsive vent. The predicted rupture range for this vehicle was 190o Âą 20 psia. During venting, unexpected motion was inparted to the spacecraft which disrupted the motion established for the passive thermal control mode. The vent tube for the supercritical helium tank is ported on two sides by diametrically opposed Oval-shaped holes. It was originally believed that the escaping gas would exit these holes at 9o degrees to the tube axis such that no net thrust is produced. However, the pressure distribution in the tube is such that the two gas plumes have an included angle less than 180 degrees and probably closer to 9o degrees. Therefore, the component of the gas flow along the axis of the vent tube produces a net thrust in the opposite direction which tends to induce a slight roll rate to the vehicle. Since +------ + +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +a. Some of the radioluminescent disks were broken. b. The apex cover was broken on the extravehicular handle side. c. The docking ring was burned and broken. d. The right--hand roll thruster was blistered. e. A yellowish/tan film existed on the outside of the hatch window, left and right rendezvous windows, and the right-hand window. f. The interior surfaces of the command module were very damp and cold, assumed to be condensation; there was no pooling of water on the floor. . Water samples could not be taken from the spacecraft tanks (discussed in section 5.8). h. The postlanding ventilation exhaust valve was open and the inlet valve was closed; the postlanding ventilation valve unlock handle was apparently jammed between the lock and unlock positions (section 14.l.2). i. There was more and deeper heat streaking in the area of the compression and shear pads than has been normally observed. + +11.0 EXPERIMENTS + +11.1 ATMOSPHERIC ELECTRICAL PHENOMENA +------ + +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:40:08 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 552...', 'Result 1:\nWhite Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat r...', 'Result 1:\nAll mechanical systems functioned properly. One mechanical anomaly, however, ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...'] +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Positions are extrapolated. + +The ll photographs have been normalized so that the earth is the same size in all frames. Frames 8590 and 8591 have been further enlarged. By viewing these two frames under a stereoscope, pronounced apparent relief is seen in the cloud patterns. The relief is so pronounced, in fact, that it cannot be attributed solely to height differences of clouds. It appears to result, in part, from the relative horizontal motion in the cloud fields; that is, clouds moving in the same direction as the spacecraft appear farther away than those moving in the direction opposite that of the spacecraftīŧŒ + +11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT +------ +Result 2: +according to ground instructions. +------ + +2025-04-03 at 19:40:08 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nPositions are extrapolated.\n\nThe ll photographs have been normalized so that ...'] +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ +Result 2: +The first reported electronic contact by the recovery forces was through S-band contact by Samoa Rescue 4. A visual sighting report by the Recovery helicopter was received and was followed shortly thereafter by aquisition of the recovery beacon signal by the Recovery, Photo, and Swim l helicopters. Fuel dump was noted and voice contact was made with the descending spacecraft, although no latitude and longitude data were received. The command module landed at 1807 G.m.t. and remained in the stable l flotation attitude. The flashing light was operating and the infiation of the uprighting system commenced about l0 minutes subsequent to landing. +------ + +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The first reported electronic contact by the recovery forces was through S-band contact by Samoa Rescue 4. A visual sighting report by the Recovery helicopter was received and was followed shortly thereafter by aquisition of the recovery beacon signal by the Recovery, Photo, and Swim l helicopters. Fuel dump was noted and voice contact was made with the descending spacecraft, although no latitude and longitude data were received. The command module landed at 1807 G.m.t. and remained in the stable l flotation attitude. The flashing light was operating and the infiation of the uprighting system commenced about l0 minutes subsequent to landing. +------ + +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 19:40:08 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nTable 5.6-I is a summary of gyro drift measurements deduced from inflight ali...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe transearth injection maneuver was performed on time, and the transearth c...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nEvent Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual con...', 'Result 1:\nThe Mission Control Center and the Manned Space Flight Network provided excel...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...'] +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 19:40:08 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 552...', 'Result 1:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 552...', 'Result 1:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 552...'] +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +14.2.2。 +------ + +2025-04-03 at 19:40:08 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 552...'] +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +14.2.2。 +------ + +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 19:40:08 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 552...', 'Result 1:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 552...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...'] +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:40:08 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.367, Perfect scores: 2/6 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.00 +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 2/6 +2025-04-03 at 19:40:08 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:40:08 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.661, Max reward: 0.881 +2025-04-03 at 19:40:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:40:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: project management objectives detailed objectives +2025-04-03 at 19:40:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 19:40:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the four primary objectives in general systems and organization? +2025-04-03 at 19:40:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-03 at 19:40:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: project management objectives types +2025-04-03 at 19:40:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 19:40:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: four primary objectives in project management +2025-04-03 at 19:40:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +TABLE 4-II.- TRAJECTORY PARAMETERS +------ + +2025-04-03 at 19:40:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what are the four primary objectives and how many detailed objectives can be derived from them +2025-04-03 at 19:40:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 19:40:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:40:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 detailed objectives +2025-04-03 at 19:40:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:40:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How many detailed objectives were derived from the four primary objectives for the lunar surface mission? +2025-04-03 at 19:40:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 19:40:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar Apollo mission aborted objectives +2025-04-03 at 19:40:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-03 at 19:40:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 11 mission primary and detailed objectives +2025-04-03 at 19:40:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:40:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:40:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 detailed objectives +2025-04-03 at 19:40:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:40:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Break down detailed objectives into the number of specific objectives +2025-04-03 at 19:40:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +------ + +2025-04-03 at 19:40:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apolo 13 mission abort objectives +2025-04-03 at 19:40:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-03 at 19:40:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 mission detailed objectives +2025-04-03 at 19:40:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:40:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:40:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 training secondary objectives +2025-04-03 at 19:40:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:40:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: break down the thirteen detailed objectives +2025-04-03 at 19:40:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 19:40:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 circumlunar return trajectory +2025-04-03 at 19:40:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:40:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:40:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: describe the four primary objectives +2025-04-03 at 19:40:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 19:40:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 power management return to earth +2025-04-03 at 19:40:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:40:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:40:29 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:40:29 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:40:29 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, True, False, True, False] +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_correctness:82 - Student lengths: [386, 554, 712, 753, 123, 514] +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_correctness:84 - Average student length: 507.00 +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 8.00 +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_correctness:86 - Length ratio: 63.38 +2025-04-03 at 19:40:29 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:40:29 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.617 Âą 0.344 +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 3.00 Âą 1.91 +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [4, 5, 5, 3, 1, 0] +2025-04-03 at 19:40:29 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +TABLE 4-II.- TRAJECTORY PARAMETERS +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +2025-04-03 at 19:40:29 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 5.0 +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.833 +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:40:29 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.233, Perfect scores: 0/6 +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.00 +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:40:29 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:40:29 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.374, Max reward: 0.618 +2025-04-03 at 19:40:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:40:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-II Engine 5 shutdown time after S-II staging +2025-04-03 at 19:40:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-03 at 19:40:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 Saturn V engine 1 shutdown time +2025-04-03 at 19:40:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 19:40:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Space Shuttle engine shutdown time after S-II staging" +2025-04-03 at 19:40:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 19:40:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:40:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: difference between lunar module engine shutdown and Apollo 13 engine 5 shutdown +2025-04-03 at 19:40:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-03 at 19:40:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 engine 5 shutdown time +2025-04-03 at 19:40:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 19:40:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:40:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module engine shutdown Apollo l3 +2025-04-03 at 19:40:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 19:40:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:40:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 engine 5 shutdown event +2025-04-03 at 19:40:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-03 at 19:40:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:40:42 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:40:42 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:40:42 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, True, True, False] +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_correctness:82 - Student lengths: [267, 200, 680, 344, 305, 213] +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [9, 9, 9, 9, 9, 9] +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_correctness:84 - Average student length: 334.83 +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 9.00 +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_correctness:86 - Length ratio: 37.20 +2025-04-03 at 19:40:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:40:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.246 Âą 0.290 +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.33 Âą 1.60 +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 0, 4, 3, 1, 0] +2025-04-03 at 19:40:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 19:40:42 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 19:40:42 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ + +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +1.0 SUMMARY + +The Apollo l3 mission, planned as a lunar landing in the Fra Mauro area, was aborted because of an abrupt loss of service module cryogenic oxygen associated with a fire in one of the two tanks at approximately 56 hours. The lunar module provided the necessary support to sustain a minimum operational. condition for a safe return to earth. A circumlunar profile was executed as the most efficient means of earth return, with the lunar module providing power and life support until transfer to the command module just prior to entry. Although the mission was unsuccessful as planned, a lunar flyby and several scientific experiments were completed. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +2025-04-03 at 19:40:42 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 19:40:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.033, Perfect scores: 0/6 +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:40:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:40:42 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.244, Max reward: 0.669 +2025-04-03 at 19:40:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:40:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar flyby mission purpose +2025-04-03 at 19:40:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:40:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:40:57 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:40:57 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:40:58 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_correctness:82 - Student lengths: [169, 9109, 5, 354, 738, 708] +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [58, 58, 58, 58, 58, 58] +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_correctness:84 - Average student length: 1847.17 +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 58.00 +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_correctness:86 - Length ratio: 31.85 +2025-04-03 at 19:40:58 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_format:228 - Responses ending properly: 0/6 +2025-04-03 at 19:40:58 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.058 Âą 0.130 +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.17 Âą 0.37 +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 0, 0, 1, 0, 0] +2025-04-03 at 19:40:58 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +2025-04-03 at 19:40:58 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +2025-04-03 at 19:40:58 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +2025-04-03 at 19:40:58 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER īŋŊīŋŊ¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +2025-04-03 at 19:40:58 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +2025-04-03 at 19:40:58 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 1.0 +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.167 +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:40:58 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:40:58 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:40:58 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.033, Max reward: 0.200 +2025-04-03 at 19:40:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:41:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: military electric field monitoring equipment locations +2025-04-03 at 19:41:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ +Result 2: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ + +2025-04-03 at 19:41:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electric field meters installed cape canaveral air force station +2025-04-03 at 19:41:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ +Result 2: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ + +2025-04-03 at 19:41:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:41:05 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:41:05 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:41:06 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, True, False] +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_correctness:82 - Student lengths: [382, 393, 350, 344, 443, 1278] +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_correctness:84 - Average student length: 531.67 +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 4.00 +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_correctness:86 - Length ratio: 132.92 +2025-04-03 at 19:41:06 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:41:06 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.096 Âą 0.137 +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.83 Âą 1.21 +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 2/6 +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 0, 0, 0, 3, 2] +2025-04-03 at 19:41:06 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 19:41:06 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 19:41:06 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 19:41:06 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 19:41:06 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ +Result 2: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ + +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ +Result 2: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ + +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:41:06 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:41:06 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:41:06 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.202, Max reward: 0.620 +2025-04-03 at 19:41:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:41:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 third midcourse correction ignition cutoff +2025-04-03 at 19:41:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:41:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: third midcourse correction ignition cutoff time +2025-04-03 at 19:41:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 19:41:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ifference of ignition cutoff in various space missions +2025-04-03 at 19:41:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 19:41:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Guidance Computer fourth and fifteenth engine commands +2025-04-03 at 19:41:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:41:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Skylab third midcourse correction ignition cutoff date +2025-04-03 at 19:41:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +159.70E 159.56E 188 371.38 188 393.19 3 065.8 3 093.2 79.364 79.934 115.464 116.54 Transearth phase Transearth injection Ignition Cutoff Thirdmidcourse correction Moon Moon 79:27 :39.0 79:32:02.8 3.73N 3.62N 65.46E 64.77E 5 465.26 5 658.68 4 547.7 5 020.2 72.645 64.784 -116.308 -117.886 Ignition Earth Earth 105:18:28.0 105:18:42.0 19.63N 19.50N 136.84W 136.90W 152 224.32 152 215.52 4 457.8 4456.6 -79.673 -79.765 114.134 114.242 Fourthmidcourse correction Ignition Cutofr Earth Earth 137:39:51.5 137:40:13.0 11.35N 11.34N 113.39E 113.32E 37 806.58 37 776.05 10 109.1 10 114.6 -72.369 -72.373 116.663 118.660 Service module separation Earth 138:01:48.0 10.88N 108.77E 35 694.93 10405.9 -71.941 118.824 Undocking Earth 141:30:00.2 1.23S 77.55E 11 257.48 1.7 465.9 -60.548 120.621 Entry interface Earth 142:40:45.7 28.23S 173.44E 65.83 36 210.6 -6.269 77.210 +------ + +2025-04-03 at 19:41:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:41:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 third midcourse correction ignition cutoff +2025-04-03 at 19:41:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 19:41:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-IVB engine cutoff time +2025-04-03 at 19:41:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 19:41:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: third midcourse correction Apollo 12 +2025-04-03 at 19:41:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:41:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:41:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 third midcourse correction ignition cutoff +2025-04-03 at 19:41:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:41:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:41:15 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:41:15 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:41:15 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, True, False, False] +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_correctness:82 - Student lengths: [95, 244, 405, 219, 364, 140] +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_correctness:84 - Average student length: 244.50 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 5.00 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_correctness:86 - Length ratio: 48.90 +2025-04-03 at 19:41:15 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:41:15 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.392 Âą 0.203 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.50 Âą 0.96 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [2, 1, 0, 2, 3, 1] +2025-04-03 at 19:41:15 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +2025-04-03 at 19:41:15 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +159.70E 159.56E 188 371.38 188 393.19 3 065.8 3 093.2 79.364 79.934 115.464 116.54 Transearth phase Transearth injection Ignition Cutoff Thirdmidcourse correction Moon Moon 79:27 :39.0 79:32:02.8 3.73N 3.62N 65.46E 64.77E 5 465.26 5 658.68 4 547.7 5 020.2 72.645 64.784 -116.308 -117.886 Ignition Earth Earth 105:18:28.0 105:18:42.0 19.63N 19.50N 136.84W 136.90W 152 224.32 152 215.52 4 457.8 4456.6 -79.673 -79.765 114.134 114.242 Fourthmidcourse correction Ignition Cutofr Earth Earth 137:39:51.5 137:40:13.0 11.35N 11.34N 113.39E 113.32E 37 806.58 37 776.05 10 109.1 10 114.6 -72.369 -72.373 116.663 118.660 Service module separation Earth 138:01:48.0 10.88N 108.77E 35 694.93 10405.9 -71.941 118.824 Undocking Earth 141:30:00.2 1.23S 77.55E 11 257.48 1.7 465.9 -60.548 120.621 Entry interface Earth 142:40:45.7 28.23S 173.44E 65.83 36 210.6 -6.269 77.210 +------ + +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 5.0 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.833 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:41:15 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.167, Perfect scores: 0/6 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.33 +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:41:15 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:41:15 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.302, Max reward: 0.763 +2025-04-03 at 19:41:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:41:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: model -d negative field value site 6 after 40 seconds +2025-04-03 at 19:41:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-03 at 19:41:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:41:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo mission site 6 temporal behavior magnetic field instability +2025-04-03 at 19:41:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 19:41:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:41:23 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:41:23 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:41:23 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 1/6 answers correct +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_correctness:82 - Student lengths: [487, 367, 225, 259, 472, 502] +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [16, 16, 16, 16, 16, 16] +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_correctness:84 - Average student length: 385.33 +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 16.00 +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_correctness:86 - Length ratio: 24.08 +2025-04-03 at 19:41:23 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:41:23 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.083 Âą 0.186 +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.33 Âą 0.75 +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [2, 0, 0, 0, 0, 0] +2025-04-03 at 19:41:23 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo l3 launch. + + + +Figure ll.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 19:41:23 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 19:41:23 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 19:41:23 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 19:41:23 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 19:41:23 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 1.0 +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.167 +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:41:23 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.067, Perfect scores: 0/6 +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.17 +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:41:23 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:41:23 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.130, Max reward: 0.781 +2025-04-03 at 19:41:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:41:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Lunar Module crew support duration +2025-04-03 at 19:41:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:41:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Four lunar day endurance mission duration limit +2025-04-03 at 19:41:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 19:41:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:41:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission crew support duration +2025-04-03 at 19:41:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 19:41:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module lunar landing power consumption and crew support duration +2025-04-03 at 19:41:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 19:41:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:41:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module power support hours after lunar module power-up +2025-04-03 at 19:41:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 19:41:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:41:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module crew support hours after second descent propulsion firing +2025-04-03 at 19:41:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 19:41:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:41:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module crew survival hours after second descent and trans Earth injection maneuver +2025-04-03 at 19:41:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 19:41:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:41:36 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:41:36 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:41:36 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, False, False, False, False] +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_correctness:82 - Student lengths: [428, 301, 206, 396, 371, 793] +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_correctness:84 - Average student length: 415.83 +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 8.00 +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_correctness:86 - Length ratio: 51.98 +2025-04-03 at 19:41:36 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:41:36 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.242 Âą 0.366 +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.17 Âą 1.86 +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 2, 0, 0, 0, 5] +2025-04-03 at 19:41:36 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +2025-04-03 at 19:41:36 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 19:41:36 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...'] +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +2025-04-03 at 19:41:36 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +2025-04-03 at 19:41:36 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +2025-04-03 at 19:41:36 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 19:41:36 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...', 'Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...'] +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:41:36 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.67 +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:41:36 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:41:36 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.197, Max reward: 0.691 +2025-04-03 at 19:41:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:41:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How far away from the line of sight to the ground station did the antenna boresight axis point due to the difference between the two sets of angles? +2025-04-03 at 19:41:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 19:41:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: + +How do geostationary satellites achieve their orbital position in relation to Earth, and what does it take to maintain this position? + + +2025-04-03 at 19:41:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-03 at 19:41:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "calculate difference in line-of-sight distance tangent angle difference +2025-04-03 at 19:41:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 ArcturugīŧŒ36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 19:41:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:41:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Iwo Jima navigation fix satellite altitude precision satellite navigation system angular distance +2025-04-03 at 19:41:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 19:41:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: calculate angle in degrees 3.263 - 0.385 +2025-04-03 at 19:41:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +-1.2 40.4 +0.4 +------ + +2025-04-03 at 19:41:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:41:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Iwo Jima orbital position satellite navigation accuracy angular distance +2025-04-03 at 19:41:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 19:41:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:41:49 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:41:49 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:41:49 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1143, 597, 466, 311, 688, 509] +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_correctness:84 - Average student length: 619.00 +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_correctness:86 - Length ratio: 61.90 +2025-04-03 at 19:41:49 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:41:49 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.250 Âą 0.265 +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.00 Âą 1.15 +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [3, 1, 0, 2, 0, 0] +2025-04-03 at 19:41:49 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-03 at 19:41:49 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nEarth-centered inertial coordinates. Bystem. **Change in velocity showm in bo...'] +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 19:41:49 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 ArcturugīŧŒ36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +-1.2 40.4 +0.4 +------ + +2025-04-03 at 19:41:49 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nTime hr:min Optian code Star used Ster angle aifference, deg Gyro torquing an...', 'Result 1:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 552...'] +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 19:41:49 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 19:41:49 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 1.0 +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.167 +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:41:49 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.033, Perfect scores: 0/6 +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:41:49 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:41:49 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.276, Max reward: 0.824 +2025-04-03 at 19:41:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:41:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module pilot laboratory studies +2025-04-03 at 19:41:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:41:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module Pilot laboratory studies outcome +2025-04-03 at 19:41:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:41:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 Command Module Pilot laboratory studies" +2025-04-03 at 19:41:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:41:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What were the results of the flight tests for the Gemini 4 Command Module Pilot laboratory studies? +2025-04-03 at 19:41:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:41:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:41:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 10 Stroking Test Analysis December 1969 +2025-04-03 at 19:41:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-03 at 19:41:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: results do not confirm Laboratory studies on Command Module Pilot +2025-04-03 at 19:41:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +------ +Result 2: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ + +2025-04-03 at 19:41:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:41:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Stroking Test Analysis +2025-04-03 at 19:41:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +The Apollo l3 space vehicle was launched from pad A of complex 39, Kennedy Space Center, Florida. Except for the high-amplitude, lowfrequency oscillations which resulted in premature cutoff of the S-II center engine, the basic performance of the launch vehicle was normal. Despite the anomaly, all launch vehicle objectives were achieved, as discussed in reference 2. In addition, the S-IvB lunar impact experiment was accomplished, as discussed in section ll.3. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:41:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: textual reference to laboratory studies on primary Command Module Pilot +2025-04-03 at 19:41:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +------ + +2025-04-03 at 19:41:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:42:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module Pilot Laboratory Analysis Apollo 13 +2025-04-03 at 19:42:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:42:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: No further information found +2025-04-03 at 19:42:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +The first reported electronic contact by the recovery forces was through S-band contact by Samoa Rescue 4. A visual sighting report by the Recovery helicopter was received and was followed shortly thereafter by aquisition of the recovery beacon signal by the Recovery, Photo, and Swim l helicopters. Fuel dump was noted and voice contact was made with the descending spacecraft, although no latitude and longitude data were received. The command module landed at 1807 G.m.t. and remained in the stable l flotation attitude. The flashing light was operating and the infiation of the uprighting system commenced about l0 minutes subsequent to landing. +------ + +2025-04-03 at 19:42:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:42:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Spaceworth Farms Command Module Laboratory Testing +2025-04-03 at 19:42:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:42:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ĐēĐžĐŧĐ°ĐŊĐ´ module landing procedure +2025-04-03 at 19:42:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:42:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:42:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Command Module Pilot Laboratory Test Results +2025-04-03 at 19:42:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:42:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:42:07 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:42:07 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:42:07 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, True, True, True] +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_correctness:82 - Student lengths: [145, 658, 280, 863, 293, 1412] +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [22, 22, 22, 22, 22, 22] +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_correctness:84 - Average student length: 608.50 +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 22.00 +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_correctness:86 - Length ratio: 27.66 +2025-04-03 at 19:42:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 19:42:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.433 Âą 0.392 +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.17 Âą 2.41 +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 1, 0, 6, 5, 1] +2025-04-03 at 19:42:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +2025-04-03 at 19:42:07 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:42:07 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe performance of the command and service module systems is discussed in thi...'] +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +2025-04-03 at 19:42:07 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +The Apollo l3 space vehicle was launched from pad A of complex 39, Kennedy Space Center, Florida. Except for the high-amplitude, lowfrequency oscillations which resulted in premature cutoff of the S-II center engine, the basic performance of the launch vehicle was normal. Despite the anomaly, all launch vehicle objectives were achieved, as discussed in reference 2. In addition, the S-IvB lunar impact experiment was accomplished, as discussed in section ll.3. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +------ +Result 2: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ + +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +------ + +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +The first reported electronic contact by the recovery forces was through S-band contact by Samoa Rescue 4. A visual sighting report by the Recovery helicopter was received and was followed shortly thereafter by aquisition of the recovery beacon signal by the Recovery, Photo, and Swim l helicopters. Fuel dump was noted and voice contact was made with the descending spacecraft, although no latitude and longitude data were received. The command module landed at 1807 G.m.t. and remained in the stable l flotation attitude. The flashing light was operating and the infiation of the uprighting system commenced about l0 minutes subsequent to landing. +------ + +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:42:07 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...'] +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:42:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.167, Perfect scores: 1/6 +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.83 +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 19:42:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:42:07 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.282, Max reward: 0.783 +2025-04-03 at 19:42:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:42:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space agency astronaut physical examination procedures +2025-04-03 at 19:42:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 19:42:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: physical examinations astronauts +2025-04-03 at 19:42:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 19:42:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:42:13 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:42:13 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:42:13 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, True, False] +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_correctness:82 - Student lengths: [238, 288, 428, 309, 440, 487] +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [25, 25, 25, 25, 25, 25] +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_correctness:84 - Average student length: 365.00 +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 25.00 +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_correctness:86 - Length ratio: 14.60 +2025-04-03 at 19:42:13 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:42:13 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.117 Âą 0.165 +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.33 Âą 0.47 +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 0, 0, 0, 1, 1] +2025-04-03 at 19:42:13 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +2025-04-03 at 19:42:13 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +2025-04-03 at 19:42:13 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +2025-04-03 at 19:42:13 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +2025-04-03 at 19:42:13 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 19:42:13 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nPostflight physical examinations were conducted immediately after recovery. T...'] +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 19:42:13 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nPostflight physical examinations were conducted immediately after recovery. T...'] +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:42:13 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 19:42:13 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:42:13 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.067, Max reward: 0.200 +2025-04-03 at 19:42:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:42:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Voyager 1 pericynthion altitude +2025-04-03 at 19:42:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:42:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pericynthion space mission planned orbital altitude +2025-04-03 at 19:42:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 19:42:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pericythion Voyager 1 maneuver orbital altitude +2025-04-03 at 19:42:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:42:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Orbital insertion pericynthion planning for Voyager 1 +2025-04-03 at 19:42:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 19:42:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the definition of pericynthion and how is it associated with orbital maneuvers in space? +2025-04-03 at 19:42:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 19:42:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:42:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Voyager 1 pericynthion altitude +2025-04-03 at 19:42:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:42:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo S-IVB translunar injection pericynthion orbit +2025-04-03 at 19:42:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 19:42:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pericynthion Voyager 1 altitude +2025-04-03 at 19:42:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 19:42:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Transearth injection Voyager 1 orbital altitude +2025-04-03 at 19:42:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:42:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:42:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Voyager 1 pericynthion location +2025-04-03 at 19:42:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 19:42:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB translunar injection velocity 23.2 ft/sec +2025-04-03 at 19:42:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:42:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pericynthion Voyager 1 post-midcourse correction altitude +2025-04-03 at 19:42:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:42:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:42:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Voyager 1 pericynthion altitude +2025-04-03 at 19:42:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:42:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pericynthion altitude of 415.8 miles midcourse maneuver +2025-04-03 at 19:42:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 19:42:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pericynthion altitude Voyager 1 w/o free return +2025-04-03 at 19:42:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:42:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:42:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pericynthion open flyby Voyager 1 +2025-04-03 at 19:42:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:42:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB free return trajectory +2025-04-03 at 19:42:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 19:42:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:42:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Voyager 1 pericynthion departure from Jupiter's orbit +2025-04-03 at 19:42:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:42:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB impact location 2.4 degrees south latitude 27.9 degrees west longitude +2025-04-03 at 19:42:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 19:42:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:42:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pericynthion of Voyager 1 +2025-04-03 at 19:42:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 19:42:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB engine mixture ratio shifts acceleration +2025-04-03 at 19:42:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 19:42:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:42:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB insertion mission control orbit deviations +2025-04-03 at 19:42:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 19:42:42 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:42:42 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:42:42 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, True] +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1879, 134, 1336, 291, 343, 216] +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_correctness:84 - Average student length: 699.83 +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 8.00 +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_correctness:86 - Length ratio: 87.48 +2025-04-03 at 19:42:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:42:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.433 Âą 0.237 +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 4.67 Âą 4.38 +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 2/6 +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [9, 0, 12, 4, 2, 1] +2025-04-03 at 19:42:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 19:42:42 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe accuracy of the first midcourse correction (table 4-Il), which placed the...', 'Result 1:\nThe accuracy of the first midcourse correction (table 4-Il), which placed the...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe accuracy of the first midcourse correction (table 4-Il), which placed the...', 'Result 1:\nThe accuracy of the first midcourse correction (table 4-Il), which placed the...', 'Result 1:\nAt approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to r...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...'] +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 19:42:42 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:42:42 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:42:43 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:42:43 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-03 at 19:42:43 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:42:43 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.17 +2025-04-03 at 19:42:43 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:42:43 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:42:43 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.370, Max reward: 0.765 +2025-04-03 at 19:42:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:42:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: window shade material stitching +2025-04-03 at 19:42:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:42:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"window shade fabric stitching materials +2025-04-03 at 19:42:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:42:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:42:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: window shade stitching material alternatives +2025-04-03 at 19:42:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:42:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"beta cloth window shade stitching material +2025-04-03 at 19:42:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:42:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:42:51 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:42:51 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:42:51 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, True, False, False] +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_correctness:82 - Student lengths: [484, 514, 133, 201, 345, 262] +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [12, 12, 12, 12, 12, 12] +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_correctness:84 - Average student length: 323.17 +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 12.00 +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_correctness:86 - Length ratio: 26.93 +2025-04-03 at 19:42:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:42:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.167 Âą 0.236 +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.67 Âą 0.94 +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 0, 2, 2, 0, 0] +2025-04-03 at 19:42:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 19:42:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 19:42:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 19:42:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 19:42:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:42:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:42:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:42:51 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.173, Max reward: 0.544 +2025-04-03 at 19:42:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:42:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: throttle position solenoid pressurization isolation 737 +2025-04-03 at 19:42:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ + +2025-04-03 at 19:42:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: solenoid, throttle position, isolation, + + +2025-04-03 at 19:42:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 19:42:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is throttle position of pressurization isolation solenoid in aircraft +2025-04-03 at 19:42:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 19:42:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: throttle position solenoid pressurization aircraft +2025-04-03 at 19:42:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 19:42:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: engine throttle position hydraulic pressurization isolation solenoid +2025-04-03 at 19:42:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +accumulator in the center-engine liquid oxygen line is being incorporated on future vehicles to decouple the line from the crossbeam, and therefore suppress any vibration amplitudes. Addition of a vibration detection system which would monitor structural response in the l4-to-20 Hz range and initiate engine cutoff if vibrations approach a dangerous level. is also under investigation as a backup. +------ + +2025-04-03 at 19:42:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressurization isolation solenoid throttle position +2025-04-03 at 19:42:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 19:42:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:42:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: throttle position, midcourse correction, descent phase, pressurization isolation Solenoid closure + + +2025-04-03 at 19:42:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:42:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressurization isolation solenoid engine shutdown +2025-04-03 at 19:42:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 19:42:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:43:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Engineering Control D helium isolation valve +2025-04-03 at 19:43:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 19:43:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:43:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure isolation solenoid throttle position closed +2025-04-03 at 19:43:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ +Result 2: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ + +2025-04-03 at 19:43:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:43:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: throttle position pressurization isolation solenoid closed 15 seconds before engine shutdown +2025-04-03 at 19:43:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:43:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:43:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: throttle position pressurization isolation solenoid closed 15 seconds before shutdown at minimum throttle +2025-04-03 at 19:43:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 19:43:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:43:06 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:43:06 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:43:06 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, False, False, False, True] +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_correctness:82 - Student lengths: [224, 925, 380, 191, 302, 260] +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [25, 25, 25, 25, 25, 25] +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_correctness:84 - Average student length: 380.33 +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 25.00 +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_correctness:86 - Length ratio: 15.21 +2025-04-03 at 19:43:06 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:43:06 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.471 Âą 0.222 +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.33 Âą 1.80 +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [1, 2, 1, 1, 3, 6] +2025-04-03 at 19:43:06 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ + +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +accumulator in the center-engine liquid oxygen line is being incorporated on future vehicles to decouple the line from the crossbeam, and therefore suppress any vibration amplitudes. Addition of a vibration detection system which would monitor structural response in the l4-to-20 Hz range and initiate engine cutoff if vibrations approach a dangerous level. is also under investigation as a backup. +------ + +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ +Result 2: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ + +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 6.0 +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 1.000 +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:43:06 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.033, Perfect scores: 0/6 +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:43:06 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:43:06 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.346, Max reward: 0.634 +2025-04-03 at 19:43:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:43:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was initial current consumption of vehicle before second descent propulsion system firing +2025-04-03 at 19:43:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:43:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space engineering curiosity rover second descent propulsion +2025-04-03 at 19:43:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ +Result 2: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ + +2025-04-03 at 19:43:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar module ascent engine consumption +2025-04-03 at 19:43:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:43:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vehicle current consumption electric second descent under propulsion system +2025-04-03 at 19:43:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-03 at 19:43:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:43:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vehicle initial current consumption of descent propulsion system firing system +2025-04-03 at 19:43:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:43:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the efficiency of a descent propulsion system +2025-04-03 at 19:43:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +With the exception of supercritical helium system performance, descent propulsion system operation, including engine starts and throttle response, was normal. +------ + +2025-04-03 at 19:43:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Falcon 9 descent 60-ampere current surge +2025-04-03 at 19:43:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 19:43:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:43:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: descent propulsion system pre-third firing initial current consumption +2025-04-03 at 19:43:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 19:43:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 descent powered-down configuration +2025-04-03 at 19:43:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:43:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:43:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: initial descent propulsion system current draw before full throttling +2025-04-03 at 19:43:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 19:43:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 powered-down configuration second descent propulsion firing +2025-04-03 at 19:43:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:43:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:43:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: descent propulsion system current consumption specific during first transearth midcourse correction +2025-04-03 at 19:43:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:43:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar landing descent phase control systems +2025-04-03 at 19:43:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:43:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:43:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first transearth midcourse descent propulsion system power draw +2025-04-03 at 19:43:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:43:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 transearth injection navigation computers +2025-04-03 at 19:43:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:43:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:43:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: initial descent propulsion system power consumption during transearth injection maneuver +2025-04-03 at 19:43:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:43:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:43:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: power consumption during transearth injection maneuver +2025-04-03 at 19:43:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ + +2025-04-03 at 19:43:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:43:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection maneuver power consumption during descent propulsion system +2025-04-03 at 19:43:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:43:26 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:43:26 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:43:26 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, True] +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_correctness:82 - Student lengths: [341, 2017, 1247, 397, 129, 510] +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_correctness:84 - Average student length: 773.50 +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_correctness:86 - Length ratio: 77.35 +2025-04-03 at 19:43:26 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:43:26 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.458 Âą 0.391 +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 3.00 Âą 3.37 +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 9, 0, 2, 1, 6] +2025-04-03 at 19:43:26 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +2025-04-03 at 19:43:26 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ + +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:43:26 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nService module.- At the time the system was powered down, reaction control sy...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...'] +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +2025-04-03 at 19:43:26 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ +Result 2: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ + +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +With the exception of supercritical helium system performance, descent propulsion system operation, including engine starts and throttle response, was normal. +------ + +2025-04-03 at 19:43:26 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nAt approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to r...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...'] +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:43:26 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe data show that descent battery 2 experienced at least a 60-ampere current...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...'] +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 1.0 +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.167 +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:43:26 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.133, Perfect scores: 0/6 +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.17 +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:43:26 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:43:26 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.375, Max reward: 0.780 +2025-04-03 at 19:43:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:43:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "spacecraft lightning strike detection event 5" +2025-04-03 at 19:43:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 19:43:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft launch lightning detection sferics detectors site 5 +2025-04-03 at 19:43:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 19:43:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: site 5 field-change and sferics detectors lightning-like discharge during launch +2025-04-03 at 19:43:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 19:43:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:43:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reliable lightning detection capabilities field-change and sferics detectors electric field meters +2025-04-03 at 19:43:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ + +2025-04-03 at 19:43:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:43:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reliability of field-change and sferics detectors in detecting lightning during launch +2025-04-03 at 19:43:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ + +2025-04-03 at 19:43:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:43:38 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:43:38 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:43:38 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, True, False, False, True] +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_correctness:82 - Student lengths: [460, 378, 438, 234, 359, 696] +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [2, 2, 2, 2, 2, 2] +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_correctness:84 - Average student length: 427.50 +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 2.00 +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_correctness:86 - Length ratio: 213.75 +2025-04-03 at 19:43:38 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 19:43:38 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.225 Âą 0.246 +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.83 Âą 1.07 +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 1, 3, 0, 0, 1] +2025-04-03 at 19:43:38 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 19:43:38 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ + +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ + +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 19:43:38 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 19:43:38 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 19:43:38 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.167, Perfect scores: 0/6 +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.17 +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 19:43:38 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:43:38 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.184, Max reward: 0.706 +2025-04-03 at 19:43:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:43:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: csm-109 manufacture history apollo program ksc +2025-04-03 at 19:43:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:43:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo commander's service module history operations location +2025-04-03 at 19:43:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:43:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "CSM 109 manufacturing facility operations history" +2025-04-03 at 19:43:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the periods when it was activated, the command module environmental control system performed normally. From the time of powering dowm at approximately 58 hours until reactivation approximately l-1/2 hours before entry, environmental control for the interconnected cabins was maintained using lumar module equipment. Two anomalies associated with the environmental control instrumentation occurred and are discussed in sections 14.l.8 and l4.l.9. An additional discrepancy, noted after landing and discussed in section l0.3, was the position of the inlet postlanding ventilation valve at the time of recovery. This discrepancy is discussed in section 14.l.2. +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 19:43:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:43:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo12 csm109 facility +2025-04-03 at 19:43:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:43:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo CSM 109 post-flight operations location +2025-04-03 at 19:43:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:43:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:43:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 10 csm manufacturer facility +2025-04-03 at 19:43:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:43:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:43:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ksc apollo10 csm operations documentation +2025-04-03 at 19:43:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:43:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:43:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: marshall space flight center csm 10 +2025-04-03 at 19:43:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:43:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:43:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: marshall space flight center apollo 10 +2025-04-03 at 19:43:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:43:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:43:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: marshall space flight center csm 109 +2025-04-03 at 19:43:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:43:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:43:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: marshall space flight center MSC-02680 +2025-04-03 at 19:43:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:43:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:43:56 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:43:56 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:43:56 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, True, False, False] +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_correctness:82 - Student lengths: [356, 1057, 432, 370, 336, 478] +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [18, 18, 18, 18, 18, 18] +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_correctness:84 - Average student length: 504.83 +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 18.00 +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_correctness:86 - Length ratio: 28.05 +2025-04-03 at 19:43:56 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:43:56 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.300 Âą 0.350 +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.83 Âą 2.85 +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 8, 0, 2, 0, 1] +2025-04-03 at 19:43:56 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 19:43:56 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:43:56 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ["Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...', 'Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...'] +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 19:43:56 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:43:56 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ["Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...'] +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 19:43:56 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases. + +The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2. + +The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4. + + + +Figure B-l.- Checkout flow for command and service modules at contractor's facility. + +NASA-S-70-5867 + + + +Figure B-2.- Command and service module checkout history at Kennedy Space Center. + +NASA-S-70-5868 + + + +Figure B-3.- Checkout flow for lunar module at contractor's facility. + + + +Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center. +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During the periods when it was activated, the command module environmental control system performed normally. From the time of powering dowm at approximately 58 hours until reactivation approximately l-1/2 hours before entry, environmental control for the interconnected cabins was maintained using lumar module equipment. Two anomalies associated with the environmental control instrumentation occurred and are discussed in sections 14.l.8 and l4.l.9. An additional discrepancy, noted after landing and discussed in section l0.3, was the position of the inlet postlanding ventilation valve at the time of recovery. This discrepancy is discussed in section 14.l.2. +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 19:43:56 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nDuring the periods when it was activated, the command module environmental co...'] +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:43:56 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.133, Perfect scores: 0/6 +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.17 +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 19:43:56 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:43:56 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.222, Max reward: 0.677 +2025-04-03 at 19:43:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:44:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: o-ring compression ratio breech assembly +2025-04-03 at 19:44:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what percentage of the block that is crushed when an O-ring enters the chamfer in a firearms breech assembly +2025-04-03 at 19:44:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 19:44:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the percentage of the O-ring that is crushed when it enters the chamfer in the breech assembly of a firearm? +2025-04-03 at 19:44:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 19:44:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +How much of the block of material is compressed by an o-ring entering a chamfer of a breech ? +2025-04-03 at 19:44:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:44:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: percentage of block crushed when O-ring enters chamfer in breech assembly +2025-04-03 at 19:44:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:44:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: valve O-ring compression ratio breech assembly 94% +2025-04-03 at 19:44:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ + +2025-04-03 at 19:44:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: >this along with some likelihood of unintended gas release due to O-ring displacement percentage late in the process +2025-04-03 at 19:44:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Percentage of block compression when O-ring enters chamfer +2025-04-03 at 19:44:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ammunition design specifications for percentage of block crushed by O-ring during breech assembly +2025-04-03 at 19:44:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:44:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:44:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Percentage of O-ring compression causing crushed attenuator block +2025-04-03 at 19:44:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What percentage of block is crushed when O-ring passes manifold port +2025-04-03 at 19:44:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:44:08 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:44:08 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:44:08 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, True, True, False] +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_correctness:82 - Student lengths: [326, 261, 128, 26, 275, 253] +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [2, 2, 2, 2, 2, 2] +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_correctness:84 - Average student length: 211.50 +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 2.00 +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_correctness:86 - Length ratio: 105.75 +2025-04-03 at 19:44:08 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:44:08 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.442 Âą 0.223 +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.83 Âą 1.07 +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [2, 1, 3, 3, 2, 0] +2025-04-03 at 19:44:08 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ + +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The possible causes of the gas leakage include: + +a. Out of tolerance parts - Measurement of the failed parts indi cate acceptable dimensions of the metal parts.. b. Damaged O-rings - The 2l 000-psi static proof-pressure test was successful. c. Gap in backup ring - The installation procedure specifies the backup ring may be trimmed on assembly to meet installation requirements, + + + +Figure 14-7.- Apex cover Jettison system. + +but does not specify any dimensional control over the scarf joint. Since the gap portion was burned away, a gap in the backup ring could have caused the problem. + +Material and dimensional controls and improvement of assembly procedures will minimize the possibility of gas leakage without necessitating a design change. However, to protect against the possibility of leaking gas with the existing design, a thermal barrier of polyimide + + + +Figure 14-8.- Damage from apex jettison thruster. + +NA SA-S-70-5849 + + + +Figure l4-9.- Plenum side of breech-plenum interface. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include: + +a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block. +2025-04-03 at 19:44:08 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 5.0 +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.833 +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:44:08 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.067, Perfect scores: 0/6 +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:44:08 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:44:08 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.454, Max reward: 0.735 +2025-04-03 at 19:44:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:44:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Redesign of valves in automotive or electric vehicles +2025-04-03 at 19:44:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: valve redesign Ball valve piston valve diaphragm valve cartridge valve +2025-04-03 at 19:44:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: solid-state battery valve replacement +2025-04-03 at 19:44:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 19:44:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: + +2025-04-03 at 19:44:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:44:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: automotive fuel valve redesign +2025-04-03 at 19:44:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:44:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: postlanding ventilation valve redesign +2025-04-03 at 19:44:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 19:44:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: valve redesign issues oxygen contamination +2025-04-03 at 19:44:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electrolyte bridging prevention valve design for batteries +2025-04-03 at 19:44:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ + +2025-04-03 at 19:44:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: supply valve postlanding aircraft oxygen isolations +2025-04-03 at 19:44:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 19:44:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:44:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: postlanding ventilation valve locking mechanism redesign +2025-04-03 at 19:44:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: postlanding ventilation inlet valve +2025-04-03 at 19:44:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electrolyte bridging prevention valve NASA Apollo +2025-04-03 at 19:44:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 19:44:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reverse leakage oxygen supply valve o-ring polytetrafluoroethylene coating +2025-04-03 at 19:44:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 19:44:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:44:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: supply valve locking mechanism preventative design +2025-04-03 at 19:44:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: valve locking mechanism malfunction +2025-04-03 at 19:44:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ + +2025-04-03 at 19:44:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: oxygen tank design failure prevention Apollo +2025-04-03 at 19:44:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 19:44:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:44:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: supply valve redesigned for O-ring installation +2025-04-03 at 19:44:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: postlanding vent valve functionality +2025-04-03 at 19:44:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo oxygen tank warning system +2025-04-03 at 19:44:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 19:44:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:44:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: supply valve redesigned for O-ring installation polytetrafluoroethylenecoated wires +2025-04-03 at 19:44:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: hydraulic cylinder lock mechanism Pin lock binding +2025-04-03 at 19:44:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 19:44:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure relief valve for oxygen tank +2025-04-03 at 19:44:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 19:44:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:44:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: valve lock problem partial travel +2025-04-03 at 19:44:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 19:44:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: oxygen tank shutoff valve design +2025-04-03 at 19:44:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:44:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: hydraulic valveactuator limitation +2025-04-03 at 19:44:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: valve O-ring design issues +2025-04-03 at 19:44:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:44:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: valve installation O-ring compression +2025-04-03 at 19:44:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 19:44:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: post-landing ventilation valve valve design +2025-04-03 at 19:44:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-03 at 19:44:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:44:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: O-ring design for self-compression +2025-04-03 at 19:44:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 19:44:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ventilation valve design mechanical issues +2025-04-03 at 19:44:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 19:44:32 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:44:32 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:44:32 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_correctness:82 - Student lengths: [152, 1270, 1777, 375, 178, 432] +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [12, 12, 12, 12, 12, 12] +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_correctness:84 - Average student length: 697.33 +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 12.00 +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_correctness:86 - Length ratio: 58.11 +2025-04-03 at 19:44:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:44:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.562 Âą 0.336 +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 5.33 Âą 4.46 +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [6, 10, 12, 0, 1, 3] +2025-04-03 at 19:44:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +a. With the handle extended only l/4 inch or less from the valve locked position, both plungers remained locked. b. With the handle extended from 5/16 to 3/8 inch from the valve locked position, the exhaust valve opened but the inlet valve remained closed.' This condition duplicates that of the position of the handle and the operation of the valve found on the Apollo l3 spacecraft after flight. c. When the handle was extended from 3/8 inch to full travel from the valve-locked position, both the inlet and and exhaust valves opened. + +Testing verified that application of power to the valves while the locking pins are being released will prevent the pin from being pulled to the unlock position because the drive shaft torque binds the lock pin. + +NASA-S-70-5841 + + + +Figure l4-l.- Post-landing vent valve lock. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +The ventilation valve is opened by first pulling the postlanding vent valve unlock handle. The handle is attached by a cable to two pins which mechanically lock the ventilation valves closed. Once the handle is pulled, the postlanding vent fan switch is placed to either the high or low position. This operation opens both ventilation valves and actuates the postlanding blower. The recovery forces found the switch setting to be proper, but the vent valve unlock handle was partially out instead of completely out. + +The inlet valve locking pin was not in the full open position (fig. l4-l), a condition which would keep the valve in the closed position even though both the pin and slot were measured to be within design tolerances. + +A check of the operation of the valves with different pull positions of the handle from locked to full open requires about one inch of travel and was made with the following results: +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +2025-04-03 at 19:44:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:44:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\n(section ll.3).\n------\nResult 2:\n+21.0 +3.0\n------\n...'] +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems. + +A more thorough discussion of this anomaly is presented in reference l. + +This anomaly is closed. + +14.1.2 Postlanding Vent Valve Malfunction + +During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:44:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.33 +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:44:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:44:32 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.523, Max reward: 0.795 +2025-04-03 at 19:44:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:44:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: immediate cause of tank line bursting in Titanic or similar +2025-04-03 at 19:44:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +is this damage which is believed to have caused the inflight failure in tank 2 and loss of pressure. +------ + +2025-04-03 at 19:44:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: cause of USS San Francisco tank line burst +2025-04-03 at 19:44:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +is this damage which is believed to have caused the inflight failure in tank 2 and loss of pressure. +------ + +2025-04-03 at 19:44:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reason for USS Iowa explosions +2025-04-03 at 19:44:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ + +2025-04-03 at 19:44:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the immediate cause of the tank line to burst in World War I +2025-04-03 at 19:44:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +is this damage which is believed to have caused the inflight failure in tank 2 and loss of pressure. +------ + +2025-04-03 at 19:44:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: battery explosions cause +2025-04-03 at 19:44:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Evidence indicates that battery 2 may have experienced an electrical fault of some type. The most probable condition is electrolyte leaking from one or more cells and bridging the high-voltage or low-voltage terminal to the battery case (fig. 14-17). This bridging results in water electrolysis and subsequent ignition of the hydrogen and oxygen so generated. The accompanying "explosion" would then blow off or rupture the seal of the battery lid and cause both a thump and venting of the free liquids in the battery case, resulting in "snowflakes." + +Postflight tests have shown the following: +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 19:44:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:44:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: USS San Francisco tethering wire insulation fire +2025-04-03 at 19:44:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 19:44:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Causes of USS Iowa accidents +2025-04-03 at 19:44:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ + +2025-04-03 at 19:44:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: battery malfunction tank line burst investigation +2025-04-03 at 19:44:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 19:44:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:44:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: USS San Francisco incident ignition cause detail +2025-04-03 at 19:44:42 | INFO | src.agent:run_tool_calls:118 - ℹīŋŊīŋŊīŋŊ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 19:44:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Apollo 13 battery explosion investigation cause +2025-04-03 at 19:44:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 19:44:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:44:44 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:44:44 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:44:44 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, False, False, False] +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_correctness:82 - Student lengths: [403, 641, 653, 585, 451, 59] +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [30, 30, 30, 30, 30, 30] +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_correctness:84 - Average student length: 465.33 +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 30.00 +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_correctness:86 - Length ratio: 15.51 +2025-04-03 at 19:44:44 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:44:44 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.413 Âą 0.222 +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.17 Âą 1.67 +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [1, 0, 3, 5, 1, 3] +2025-04-03 at 19:44:44 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +is this damage which is believed to have caused the inflight failure in tank 2 and loss of pressure. +------ + +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +2025-04-03 at 19:44:44 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +is this damage which is believed to have caused the inflight failure in tank 2 and loss of pressure. +------ + +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Following the anomaly, all reaction control quads except C began showing evidence of frequent engine firings. Data show that all propellant isolation valves Cn quad C, both helium isolation valves on quad D, and one helium isolation valve on quad B were shocked to the closed posi-- tior at the time of the oxygen tank pressure loss. On quad D, the regulated pressures dropped momentarily as the engines fired with the helium isolation valves closed. The crew reopened the quad D valves, and the engines functioned normally thereafter. Because the quad C propellant isolation valves are powered from bus B, which lost power, the valves could not be reopened and the quad remained inactive for the remainder of the flight. +------ + +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ + +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +is this damage which is believed to have caused the inflight failure in tank 2 and loss of pressure. +------ + +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Evidence indicates that battery 2 may have experienced an electrical fault of some type. The most probable condition is electrolyte leaking from one or more cells and bridging the high-voltage or low-voltage terminal to the battery case (fig. 14-17). This bridging results in water electrolysis and subsequent ignition of the hydrogen and oxygen so generated. The accompanying "explosion" would then blow off or rupture the seal of the battery lid and cause both a thump and venting of the free liquids in the battery case, resulting in "snowflakes." + +Postflight tests have shown the following: +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 5.0 +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.833 +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:44:44 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.233, Perfect scores: 0/6 +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.50 +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 19:44:44 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:44:44 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.367, Max reward: 0.634 +2025-04-03 at 19:44:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:44:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: definition of 13-60-8591 frame +2025-04-03 at 19:44:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:44:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: object 13-60-8591 latitude search +2025-04-03 at 19:44:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 19:44:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: latency latitude object frame 13-60-8591 +2025-04-03 at 19:44:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 ArcturugīŧŒ36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ + +2025-04-03 at 19:44:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:44:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 in degrees +2025-04-03 at 19:44:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +-1.2 40.4 +0.4 +------ + +2025-04-03 at 19:44:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Beachcraft retrieval Iwo Jima latitude +2025-04-03 at 19:44:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +After confirrning the integrity of the command module and the status of the crew, the Recovery helicopter crew attempted to recover the main parachutes with grappling hooks and flotation gear prior to their sinking. Swim l and Swim 2 helicopters arrived on scene and immediately proceeded with retrieval. Swim 2 deployed swimmers to provide flotation to the spacecraft, and Swim l deployed swimmers to retrieve the apex cover, which was located upwind of the spacecraft. The flight crew was onboard the recovery helicopter 7 minutes after they had egressed the command module, and they arrived aboard Iwo Jima at 1853 G.m.t. + +Command module retrieval took place at 2l degrees 39.l minutes south latitude and 165 degrees 20.9 minutes west longitude at 1936 G.m.t. One main parachute and the apex cover were retrieved by small boat and brought aboard. +------ + +2025-04-03 at 19:44:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: starlink satellite 13-60-8591 latitude +2025-04-03 at 19:44:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 19:44:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:44:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: asteroid 13-60-8591 latitude +2025-04-03 at 19:44:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 ArcturugīŧŒ36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ + +2025-04-03 at 19:44:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 landing site latitude +2025-04-03 at 19:44:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:44:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:44:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: asteroid 13-60-8591 latitude 40.4 +2025-04-03 at 19:44:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 ArcturugīŧŒ36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ + +2025-04-03 at 19:44:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:44:55 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:44:55 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:44:55 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_correctness:82 - Student lengths: [387, 31, 353, 718, 258, 512] +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [7, 7, 7, 7, 7, 7] +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_correctness:84 - Average student length: 376.50 +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 7.00 +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_correctness:86 - Length ratio: 53.79 +2025-04-03 at 19:44:55 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:44:55 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.325 Âą 0.336 +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.50 Âą 1.61 +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 4, 2, 0, 0, 3] +2025-04-03 at 19:44:55 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 19:44:55 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +-1.2 40.4 +0.4 +------ + +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 ArcturugīŧŒ36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ + +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 ArcturugīŧŒ36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ + +2025-04-03 at 19:44:55 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 552...', 'Result 1:\n1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 552...', 'Result 1:\nvelocity vector onto the local body-centered, horizontal plane, measured posi...', 'Result 1:\nvelocity vector onto the local body-centered, horizontal plane, measured posi...'] +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +After confirrning the integrity of the command module and the status of the crew, the Recovery helicopter crew attempted to recover the main parachutes with grappling hooks and flotation gear prior to their sinking. Swim l and Swim 2 helicopters arrived on scene and immediately proceeded with retrieval. Swim 2 deployed swimmers to provide flotation to the spacecraft, and Swim l deployed swimmers to retrieve the apex cover, which was located upwind of the spacecraft. The flight crew was onboard the recovery helicopter 7 minutes after they had egressed the command module, and they arrived aboard Iwo Jima at 1853 G.m.t. + +Command module retrieval took place at 2l degrees 39.l minutes south latitude and 165 degrees 20.9 minutes west longitude at 1936 G.m.t. One main parachute and the apex cover were retrieved by small boat and brought aboard. +------ + +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 19:44:55 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 19:44:55 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 ArcturugīŧŒ36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ + +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ +Result 2: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ + +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:44:55 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.033, Perfect scores: 0/6 +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:44:55 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:44:55 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.403, Max reward: 0.818 +2025-04-03 at 19:44:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:45:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA primary crew physical examinations before launch +2025-04-03 at 19:45:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +------ +Result 2: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ + +2025-04-03 at 19:45:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "NASA astronaut physical examination procedure before launch" +2025-04-03 at 19:45:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 19:45:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: crew physical examinations pre-launch Apollo mission +2025-04-03 at 19:45:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 19:45:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:45:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 primary crew prelaunch physical examinations +2025-04-03 at 19:45:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +------ + +2025-04-03 at 19:45:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:45:03 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:45:03 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:45:04 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, True, True, False] +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_correctness:82 - Student lengths: [269, 515, 422, 1519, 389, 481] +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [1, 1, 1, 1, 1, 1] +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_correctness:84 - Average student length: 599.17 +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 1.00 +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_correctness:86 - Length ratio: 599.17 +2025-04-03 at 19:45:04 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:45:04 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.200 Âą 0.206 +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.67 Âą 0.75 +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [2, 0, 1, 0, 1, 0] +2025-04-03 at 19:45:04 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +------ +Result 2: +Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses. + +The Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites. + +10.0 MISSION SUPPORT PERFORMANCE + +10.1 FLIGHT CONTROL +------ + +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +------ + +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +2025-04-03 at 19:45:04 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 19:45:04 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...'] +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +2025-04-03 at 19:45:04 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +2025-04-03 at 19:45:04 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:45:04 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.033, Perfect scores: 0/6 +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:45:04 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:45:04 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.155, Max reward: 0.531 +2025-04-03 at 19:45:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:45:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Site 5 device purpose +2025-04-03 at 19:45:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +according to ground instructions. +------ +Result 2: +urine trans. fer line. Also included was a lunar topographic camera, which could be installed in the command module hatch window for high resolution photography of the lunar surface from orbit. The camera provided a 4.5-inch film format and had an 18-inch focal length and image-motion compensation. The photographs would yield a resolution of approximately l2 feet and would include a l5-mile square area on the surface for each frame exposed. +------ + +2025-04-03 at 19:45:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Site 5 nuclear reactor device purpose +2025-04-03 at 19:45:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-03 at 19:45:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: site 5 nuclear experiment +2025-04-03 at 19:45:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ + +2025-04-03 at 19:45:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: site 5 carbon ablator system Apollo 13 +2025-04-03 at 19:45:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:45:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Liquid core nuclear reactor site 5" +2025-04-03 at 19:45:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:45:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:45:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: integrated opticon purpose +2025-04-03 at 19:45:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-03 at 19:45:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission system operations at Site 5 +2025-04-03 at 19:45:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:45:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission site 5 investigation +2025-04-03 at 19:45:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:45:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:45:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: operator interface device site 5 +2025-04-03 at 19:45:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:45:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission specific item at Site 5 +2025-04-03 at 19:45:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:45:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system Apollo decompression +2025-04-03 at 19:45:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:45:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:45:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: optics coupling display unit +2025-04-03 at 19:45:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 19:45:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 reaction control system evaluation +2025-04-03 at 19:45:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:45:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:45:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: site 5 space suit testing facility +2025-04-03 at 19:45:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:45:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:45:18 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:45:18 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:45:18 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, False, False, False, False] +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_correctness:82 - Student lengths: [128, 833, 279, 300, 452, 269] +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [46, 46, 46, 46, 46, 46] +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_correctness:84 - Average student length: 376.83 +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 46.00 +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_correctness:86 - Length ratio: 8.19 +2025-04-03 at 19:45:18 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:45:18 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.517 Âą 0.318 +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.33 Âą 1.80 +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [4, 3, 1, 0, 5, 1] +2025-04-03 at 19:45:18 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +according to ground instructions. +------ +Result 2: +urine trans. fer line. Also included was a lunar topographic camera, which could be installed in the command module hatch window for high resolution photography of the lunar surface from orbit. The camera provided a 4.5-inch film format and had an 18-inch focal length and image-motion compensation. The photographs would yield a resolution of approximately l2 feet and would include a l5-mile square area on the surface for each frame exposed. +------ + +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 19:45:18 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\naccording to ground instructions.\n------\nResult 2:\nurine trans. fer line. Als...', 'Result 1:\ncause for optics coupling display unit indications ofoptics movement during t...', 'Result 1:\ncause for optics coupling display unit indications ofoptics movement during t...', 'Result 1:\ncause for optics coupling display unit indications ofoptics movement during t...'] +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:45:18 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...'] +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ + +2025-04-03 at 19:45:18 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe charged particle lunar environment experiment was designed to measure the...'] +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 19:45:18 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:45:18 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...'] +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:45:18 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe field-change and sferics detectors at site 5 gave no indication of any li...'] +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:45:18 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:45:18 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:45:18 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.393, Max reward: 0.662 +2025-04-03 at 19:45:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:45:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"X-Scale factor definition electronic scale error" +2025-04-03 at 19:45:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 19:45:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: X-scale factor ppm X-ray inspection solder joints +2025-04-03 at 19:45:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transducer is a variable reluctance instrument actuated by differential pressure applied across a twisted Bourdon tube. The housing, including the cavity containing the Bourdon tube and the variable reluctance elements, is nickel plated. The Bourdon tube-variable reluctance + +assembly and the sense port fitting are soldered in place. Inspection of the failed units indicates that the flaking occurs adjacent to the solder. + +The most probable cause of the problem is poor plating adhesion to the aluminum base metal. Differential expansion between the solder and the aluminum may cause the plating to crack. Moisture from the environmental control system suit loop could then penetrate the plating, corrode the aluminum base metal and cause the plating to peel and flake. The nickel flakes could then enter the air gap of the variable reluctance elements and affect the measurement. +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-03 at 19:45:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "X-Scale factor error ppm" +2025-04-03 at 19:45:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 19:45:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: X-ray scale factor error in ppm +2025-04-03 at 19:45:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 19:45:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:45:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "X-Scale factor error ppm relation to gyrocompassing" +2025-04-03 at 19:45:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:45:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: X-Scale factor unit in ppm +2025-04-03 at 19:45:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:45:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:45:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "X-Scale factor error gyrocompassing Y-axis error magnitude" +2025-04-03 at 19:45:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:45:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: X-Scale factor error unit +2025-04-03 at 19:45:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 19:45:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:45:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "S-IvB instrument unit gyrocompassing misalignment" +2025-04-03 at 19:45:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 19:45:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: X- Scale factor error unit +2025-04-03 at 19:45:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 19:45:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:45:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "S-IvB instrument unit Y-axis error magnitude null bias drift" +2025-04-03 at 19:45:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:45:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: X- Scale factor error unit +2025-04-03 at 19:45:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 19:45:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:45:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "null bias drift acceleration drift X-Scale factor error ppm" +2025-04-03 at 19:45:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:45:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: X-Scale factor error unit +2025-04-03 at 19:45:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 19:45:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:45:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "X-Scale factor error null bias drift relation" +2025-04-03 at 19:45:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 19:45:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:45:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "X-Scale factor error acceleration drift correlation" +2025-04-03 at 19:45:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:45:37 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:45:37 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:45:37 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_correctness:82 - Student lengths: [266, 269, 449, 1697, 1435, 227] +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [1, 1, 1, 1, 1, 1] +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_correctness:84 - Average student length: 723.83 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 1.00 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_correctness:86 - Length ratio: 723.83 +2025-04-03 at 19:45:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:45:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.433 Âą 0.392 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.67 Âą 3.14 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 1, 1, 8, 6, 0] +2025-04-03 at 19:45:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +2025-04-03 at 19:45:37 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The transducer is a variable reluctance instrument actuated by differential pressure applied across a twisted Bourdon tube. The housing, including the cavity containing the Bourdon tube and the variable reluctance elements, is nickel plated. The Bourdon tube-variable reluctance + +assembly and the sense port fitting are soldered in place. Inspection of the failed units indicates that the flaking occurs adjacent to the solder. + +The most probable cause of the problem is poor plating adhesion to the aluminum base metal. Differential expansion between the solder and the aluminum may cause the plating to crack. Moisture from the environmental control system suit loop could then penetrate the plating, corrode the aluminum base metal and cause the plating to peel and flake. The nickel flakes could then enter the air gap of the variable reluctance elements and affect the measurement. +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-03 at 19:45:37 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe transducer is a variable reluctance instrument actuated by differential p...'] +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ + +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +2025-04-03 at 19:45:37 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 19:45:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.33 +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:45:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:45:37 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.067, Max reward: 0.200 +2025-04-03 at 19:45:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:45:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space x moon mission crew command module pilot +2025-04-03 at 19:45:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 19:45:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacex crew dragon commercial crew mission 2 crew swap +2025-04-03 at 19:45:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:45:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:45:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: SpaceX Crew-6 lunar module to spacecraft separation +2025-04-03 at 19:45:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:45:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "spacex crew dragon commercial crew mission 2 mission details" +2025-04-03 at 19:45:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:45:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:45:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: SpaceX Crew-6 lunar orbit insertion +2025-04-03 at 19:45:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:45:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacex crew dragon commercial crew mission 2 crew roles +2025-04-03 at 19:45:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ + +2025-04-03 at 19:45:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:45:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: SpaceX Crew-6 lunar ferry crew commanders +2025-04-03 at 19:45:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:45:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacex crew-2 mission crew rotation and cryogenic oxygen tank incident +2025-04-03 at 19:45:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ +Result 2: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ + +2025-04-03 at 19:45:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:45:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: SpaceX Crew-6 Commander Pilot +2025-04-03 at 19:45:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-03 at 19:45:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacex crew-2 abort decision making cryogenic tank failure +2025-04-03 at 19:45:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 19:45:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:45:51 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:45:51 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:45:51 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 1/6 answers correct +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, False, False] +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_correctness:82 - Student lengths: [377, 305, 396, 210, 629, 590] +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [20, 20, 20, 20, 20, 20] +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_correctness:84 - Average student length: 417.83 +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 20.00 +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_correctness:86 - Length ratio: 20.89 +2025-04-03 at 19:45:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:45:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.317 Âą 0.448 +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.67 Âą 2.36 +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [5, 5, 0, 0, 0, 0] +2025-04-03 at 19:45:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:368 - īŋŊīŋŊ Searched Chunk 1: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-03 at 19:45:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe unprecedented powered-down state of the command module required generatio...', 'Result 1:\nFollowing separation and translation, a manual pitch maneuver of 1.5 deg/sec ...', 'Result 1:\nAt approximately 105 hours, the crew performed a manual descent propulsion ma...', 'Result 1:\nCommander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., an...', 'Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...'] +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ + +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ +Result 2: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ + +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 19:45:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe structure in the forward end of the docking tunnel was reinforced to acco...', 'Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...', 'Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...', 'Result 1:\nAt approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to r...', 'Result 1:\nThe two tanks cortaining cryogenic oxygen, used for _fuel cell operation and ...'] +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +2025-04-03 at 19:45:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +2025-04-03 at 19:45:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +2025-04-03 at 19:45:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +2025-04-03 at 19:45:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:45:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.50 +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:45:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:45:51 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.240, Max reward: 0.757 +2025-04-03 at 19:45:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:45:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +computer-displayed inertial velocity "final haircut" + +2025-04-03 at 19:45:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 19:45:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:45:58 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:45:58 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:45:58 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_correctness:82 - Student lengths: [255, 201, 490, 528, 328, 247] +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [13, 13, 13, 13, 13, 13] +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_correctness:84 - Average student length: 341.50 +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 13.00 +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_correctness:86 - Length ratio: 26.27 +2025-04-03 at 19:45:58 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:45:58 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.058 Âą 0.130 +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.17 Âą 0.37 +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 0, 0, 0, 1, 0] +2025-04-03 at 19:45:58 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 19:45:58 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 19:45:58 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 19:45:58 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 19:45:58 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 19:45:58 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nEvent Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product o...'] +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 19:45:58 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:45:58 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:45:58 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:45:58 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.033, Max reward: 0.200 +2025-04-03 at 19:45:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:46:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: asa spacecraft launch august 25 1966 AS-202 SC-011 +2025-04-03 at 19:46:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Mi ssion Spacecraft Description Launch date Launch site Apollo4 SC-017 LTA-10R Supercircular entry at lunar Nov.9,1967 Kennedy Space Center, Fla. Apollo 5 LM-1 return velocity First lunar module flight Jan.22,1968 Cape Kennedy, Apollo 6 SC-020 LTA-2R Verification of closed-loop April 4, 1968 Fla. Kennedy Space Center, Fla. Apollo7 CSM 101 emergency detection system First manned flight; Oct.11īŧŒ1968 Apol1o 8 CSM 103 earth-orbital First manned lunar Dec.2l,1968 Cape Kennedy, Fla. Kennedy Space Apol1o9 CSM 104 orbital flight; first manned Saturn V launch First manned lunar Apollo 10 LM-3 module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center, Fla. CSM 106 LM-4 First lunar orbit rendezvous; low pass over lumar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 Second lunar landing Center, Fla. Apollo 13 LM-6 Nov. 14, 1969 Kennedy Space Center, Fla. +------ + +2025-04-03 at 19:46:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-202 SC-011 incident type August 25 1966 +2025-04-03 at 19:46:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ +Result 2: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ + +2025-04-03 at 19:46:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: askanasa mission as 202 sc 011 august 25 1966 +2025-04-03 at 19:46:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 19:46:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:46:03 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:46:03 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:46:03 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_correctness:82 - Student lengths: [299, 273, 588, 137, 89, 90] +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [39, 39, 39, 39, 39, 39] +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_correctness:84 - Average student length: 246.00 +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 39.00 +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_correctness:86 - Length ratio: 6.31 +2025-04-03 at 19:46:03 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:46:03 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.175 Âą 0.175 +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.50 Âą 0.50 +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 0, 1, 1, 0, 1] +2025-04-03 at 19:46:03 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 19:46:03 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 19:46:03 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Mi ssion Spacecraft Description Launch date Launch site Apollo4 SC-017 LTA-10R Supercircular entry at lunar Nov.9,1967 Kennedy Space Center, Fla. Apollo 5 LM-1 return velocity First lunar module flight Jan.22,1968 Cape Kennedy, Apollo 6 SC-020 LTA-2R Verification of closed-loop April 4, 1968 Fla. Kennedy Space Center, Fla. Apollo7 CSM 101 emergency detection system First manned flight; Oct.11īŧŒ1968 Apol1o 8 CSM 103 earth-orbital First manned lunar Dec.2l,1968 Cape Kennedy, Fla. Kennedy Space Apol1o9 CSM 104 orbital flight; first manned Saturn V launch First manned lunar Apollo 10 LM-3 module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center, Fla. CSM 106 LM-4 First lunar orbit rendezvous; low pass over lumar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 Second lunar landing Center, Fla. Apollo 13 LM-6 Nov. 14, 1969 Kennedy Space Center, Fla. +------ + +2025-04-03 at 19:46:03 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, ...'] +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ +Result 2: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ + +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 19:46:03 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 19:46:03 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nMission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad ab...'] +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 1.0 +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.167 +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:46:03 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:46:03 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:46:03 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.100, Max reward: 0.200 +2025-04-03 at 19:46:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:46:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "found on spaceship or rocket propellant gas escape aftermath" and "adverse effects of propellant gas spills" and "chemical propellant spills on equipment" and "components found after propellant gas leak" and "hazardous materials on spacecraft" and "enabled components contaminated by propellant gas" and "escapes propellant gas affected systems" +2025-04-03 at 19:46:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +a. Some of the radioluminescent disks were broken. b. The apex cover was broken on the extravehicular handle side. c. The docking ring was burned and broken. d. The right--hand roll thruster was blistered. e. A yellowish/tan film existed on the outside of the hatch window, left and right rendezvous windows, and the right-hand window. f. The interior surfaces of the command module were very damp and cold, assumed to be condensation; there was no pooling of water on the floor. . Water samples could not be taken from the spacecraft tanks (discussed in section 5.8). h. The postlanding ventilation exhaust valve was open and the inlet valve was closed; the postlanding ventilation valve unlock handle was apparently jammed between the lock and unlock positions (section 14.l.2). i. There was more and deeper heat streaking in the area of the compression and shear pads than has been normally observed. + +11.0 EXPERIMENTS + +11.1 ATMOSPHERIC ELECTRICAL PHENOMENA +------ + +2025-04-03 at 19:46:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: propellant gas contamination effects on equipment + + +2025-04-03 at 19:46:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:46:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was found on adjacent equipment due to escaped rocket propellant gas +2025-04-03 at 19:46:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:46:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is typically found on adjacent equipment due to an escaped propellant gas in a rocket or missile propulsion system? +2025-04-03 at 19:46:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:46:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:46:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was found on adjacent equipment due to escaped rocket propellant gas +2025-04-03 at 19:46:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:46:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what kind of debris was deposited on the surrounding equipment due to the escaped propellant gas from the breech plenum interface? +2025-04-03 at 19:46:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:46:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:46:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the composition of the soot deposited on adjacent equipment due to the escaped propellant gas? +2025-04-03 at 19:46:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:46:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:46:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what compounds are present in the combustion residue deposited on adjacent equipment due to escaped propellant gas? +2025-04-03 at 19:46:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 19:46:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:46:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what type of insulation material is the damaged one mentioned? +2025-04-03 at 19:46:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 19:46:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:46:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: contaminant mentioned in the information that degrades the insulating characteristics of the vacuum tank and causes overheating +2025-04-03 at 19:46:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 19:46:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:46:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: cause of the wire insulation burning +2025-04-03 at 19:46:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ + +2025-04-03 at 19:46:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:46:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: cause of the short between the temperature switch wires to ground +2025-04-03 at 19:46:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 19:46:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:46:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: short circuit location in the electrical system +2025-04-03 at 19:46:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 19:46:23 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:46:23 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:46:23 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_correctness:82 - Student lengths: [549, 1377, 550, 750, 260, 1756] +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_correctness:84 - Average student length: 873.67 +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 4.00 +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_correctness:86 - Length ratio: 218.42 +2025-04-03 at 19:46:23 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_format:228 - Responses ending properly: 0/6 +2025-04-03 at 19:46:23 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.358 Âą 0.323 +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.17 Âą 3.13 +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 1, 0, 1, 2, 9] +2025-04-03 at 19:46:23 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 19:46:23 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +a. Some of the radioluminescent disks were broken. b. The apex cover was broken on the extravehicular handle side. c. The docking ring was burned and broken. d. The right--hand roll thruster was blistered. e. A yellowish/tan film existed on the outside of the hatch window, left and right rendezvous windows, and the right-hand window. f. The interior surfaces of the command module were very damp and cold, assumed to be condensation; there was no pooling of water on the floor. . Water samples could not be taken from the spacecraft tanks (discussed in section 5.8). h. The postlanding ventilation exhaust valve was open and the inlet valve was closed; the postlanding ventilation valve unlock handle was apparently jammed between the lock and unlock positions (section 14.l.2). i. There was more and deeper heat streaking in the area of the compression and shear pads than has been normally observed. + +11.0 EXPERIMENTS + +11.1 ATMOSPHERIC ELECTRICAL PHENOMENA +------ + +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 19:46:23 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ + +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ + +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:46:23 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.200, Perfect scores: 1/6 +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.83 +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 19:46:23 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:46:23 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.178, Max reward: 0.667 +2025-04-03 at 19:46:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:46:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "NASA Mars Orbital Habitat size +2025-04-03 at 19:46:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:46:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: environmental control system habitable environment crew duration +2025-04-03 at 19:46:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +During the periods when it was activated, the command module environmental control system performed normally. From the time of powering dowm at approximately 58 hours until reactivation approximately l-1/2 hours before entry, environmental control for the interconnected cabins was maintained using lumar module equipment. Two anomalies associated with the environmental control instrumentation occurred and are discussed in sections 14.l.8 and l4.l.9. An additional discrepancy, noted after landing and discussed in section l0.3, was the position of the inlet postlanding ventilation valve at the time of recovery. This discrepancy is discussed in section 14.l.2. +------ + +2025-04-03 at 19:46:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:46:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ex habitable environment duration Apollo +2025-04-03 at 19:46:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 19:46:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 17 environmental control system +2025-04-03 at 19:46:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:46:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:46:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: hours idss Apollo crew habitable environment +2025-04-03 at 19:46:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew reported sleeping well the first 2 days of the mission. They all slept about 5-l/2 hours during the first sleep period. During the second period, the Commander, Command Module Pilot, and Lunar Module Pilot slept 5, 6, and 9 hours, respectively. The third sleep period was scheduled for 6l hours, but the oxygen tank incident at 56 hours precluded sleep by any of the crew until approximately 8o hours. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 19:46:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 17 environmental control system habitable environment duration +2025-04-03 at 19:46:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:46:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:46:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no +2025-04-03 at 19:46:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +(section ll.3). +------ + +2025-04-03 at 19:46:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 17 environmental control system crew habitation time +2025-04-03 at 19:46:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:46:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:46:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo crew recovery environment hours after landing +2025-04-03 at 19:46:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 19:46:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew habitation system duration environmental control +2025-04-03 at 19:46:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 19:46:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:46:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apolo crew recovery duration from lunar landing> +2025-04-03 at 19:46:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:46:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew habitation environment duration +2025-04-03 at 19:46:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:46:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:46:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apolo crew recovery time from lunar landing crew +2025-04-03 at 19:46:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:46:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission duration crew environment +2025-04-03 at 19:46:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:46:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:46:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew environment duration +2025-04-03 at 19:46:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:46:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:46:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew survival environment duration +2025-04-03 at 19:46:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:46:41 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:46:41 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:46:42 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, False, False, False, False] +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_correctness:82 - Student lengths: [716, 158, 191, 1984, 256, 433] +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [2, 2, 2, 2, 2, 2] +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_correctness:84 - Average student length: 623.00 +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 2.00 +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_correctness:86 - Length ratio: 311.50 +2025-04-03 at 19:46:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 19:46:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.317 Âą 0.448 +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.67 Âą 3.82 +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 7, 0, 9, 0, 0] +2025-04-03 at 19:46:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +2025-04-03 at 19:46:42 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The crew reported sleeping well the first 2 days of the mission. They all slept about 5-l/2 hours during the first sleep period. During the second period, the Commander, Command Module Pilot, and Lunar Module Pilot slept 5, 6, and 9 hours, respectively. The third sleep period was scheduled for 6l hours, but the oxygen tank incident at 56 hours precluded sleep by any of the crew until approximately 8o hours. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: ++21.0 +3.0 +------ +Result 2: +(section ll.3). +------ + +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +The lunar module was jettisoned l hour before entry, which was performed nominally using the primary guidance and navigation system. Landing occurred at 142:54:4l within sight of the recovery ship. The landing point was reported as 2l degrees 38 minutes 24 seconds south latitude and l65 degrees 2l minutes 42 seconds west longitude. The crew were retrieved and aboard the recovery ship within 45 minutes after landing。 + +2.0 INTRODUCTION + +Apollo l3 was the thirteenth in a series of missions using Apollo specification flight hardware and was to be the third lunar landing. The primary mission objective was a precise lunar landing to conduct scientific exploration of deep-rooted surface material. +------ + +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +2025-04-03 at 19:46:42 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +During the periods when it was activated, the command module environmental control system performed normally. From the time of powering dowm at approximately 58 hours until reactivation approximately l-1/2 hours before entry, environmental control for the interconnected cabins was maintained using lumar module equipment. Two anomalies associated with the environmental control instrumentation occurred and are discussed in sections 14.l.8 and l4.l.9. An additional discrepancy, noted after landing and discussed in section l0.3, was the position of the inlet postlanding ventilation valve at the time of recovery. This discrepancy is discussed in section 14.l.2. +------ + +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +2025-04-03 at 19:46:42 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +2025-04-03 at 19:46:42 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:46:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:46:42 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:46:42 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.142, Max reward: 0.692 +2025-04-03 at 19:46:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:46:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what triggered the switch to manual system? +2025-04-03 at 19:46:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +This anomaly is open. + +14.1.5 Entry Monitor System 0.05g Light Malfunction + +The entry monitor system 0.05g.light did not illuminate within 3 seconds after an 0.05g condition was sensed by the guidance system. The crew started the system manually as prescribed by switching to the backupposition. + +The entry monitor system is designed to start automatically when 0.05g is sensed by the system accelerometer. When this sensing occurs, the 0.05g light should come on, the scroll should begin to drive, and the lrange-to-go counter should begin to count down. The crew reported the light failure but were unable to verify whether or not the scroll or counter responded before the switch was manually changed to the backup mode. + +The failure had to be in the light, in the 0.05g sensing mechanism, or in the mode switch, mode switching could also have been premature. +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-03 at 19:46:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "when did ships switch to manual systems" +2025-04-03 at 19:46:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 19:46:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what triggered the switch to manual systems in workplaces during the 1970s +2025-04-03 at 19:46:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ + +2025-04-03 at 19:46:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:46:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: atlantis manual system switch anomaly +2025-04-03 at 19:46:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ +Result 2: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ + +2025-04-03 at 19:46:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "what triggered switchover from automatic to manual in space exploration" +2025-04-03 at 19:46:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 19:46:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:46:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 14 manual system switch causes +2025-04-03 at 19:46:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ +Result 2: +Performance of the guidance, navigation, and control system was normal except for two instances. Random motion observed in the sextant shaft during the zero optics mode was operationally prevented by turning off power to the optical system when not in use. This problem occurred during Apollo l2 and is thought to be caused by a buildup of contact resistance in the slip rings of the half-speed resolver in the sextant (section 14.l.3). The crew reported the 0.05g light did not illuminate as required within 3 seconds after the digital computer had indicated 0.05g. A manual procedure was therefore required to start the entry monitor system, which performed nominally throughout the remainder of entry (section 14.l.5). As a result of the aborted mission, all power was removed from the inertial platform, including heaters, for approximately 80 hours. After powering up and coarse aligning the platform to that of the lunar module, the cormmand module was guided to a successful landing within +------ + +2025-04-03 at 19:46:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "specific incident triggered Apollo switchover to manual control" +2025-04-03 at 19:46:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-03 at 19:46:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:46:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 14 sextant resolver contact resistance +2025-04-03 at 19:46:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of the guidance, navigation, and control system was normal except for two instances. Random motion observed in the sextant shaft during the zero optics mode was operationally prevented by turning off power to the optical system when not in use. This problem occurred during Apollo l2 and is thought to be caused by a buildup of contact resistance in the slip rings of the half-speed resolver in the sextant (section 14.l.3). The crew reported the 0.05g light did not illuminate as required within 3 seconds after the digital computer had indicated 0.05g. A manual procedure was therefore required to start the entry monitor system, which performed nominally throughout the remainder of entry (section 14.l.5). As a result of the aborted mission, all power was removed from the inertial platform, including heaters, for approximately 80 hours. After powering up and coarse aligning the platform to that of the lunar module, the cormmand module was guided to a successful landing within +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:46:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 thrust/translation controller assembly operations" +2025-04-03 at 19:46:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:46:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:46:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 14 landing success +2025-04-03 at 19:46:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:46:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 system malfunction manual override" +2025-04-03 at 19:46:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:46:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:46:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 hydrogen tank issues +2025-04-03 at 19:46:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 19:46:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:46:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar ascent and service module oxygen tank issues +2025-04-03 at 19:46:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 19:46:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:47:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 report section 14.5 +2025-04-03 at 19:47:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:47:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:47:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 crew mission briefing +2025-04-03 at 19:47:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:47:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:47:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 training itdp +2025-04-03 at 19:47:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:47:02 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:47:02 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:47:02 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, True, False, False] +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_correctness:82 - Student lengths: [679, 774, 1984, 627, 346, 257] +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [28, 28, 28, 28, 28, 28] +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_correctness:84 - Average student length: 777.83 +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 28.00 +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_correctness:86 - Length ratio: 27.78 +2025-04-03 at 19:47:02 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:47:02 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.375 Âą 0.425 +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.67 Âą 3.73 +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 0, 10, 5, 1, 0] +2025-04-03 at 19:47:02 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +2025-04-03 at 19:47:02 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +2025-04-03 at 19:47:02 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +This anomaly is open. + +14.1.5 Entry Monitor System 0.05g Light Malfunction + +The entry monitor system 0.05g.light did not illuminate within 3 seconds after an 0.05g condition was sensed by the guidance system. The crew started the system manually as prescribed by switching to the backupposition. + +The entry monitor system is designed to start automatically when 0.05g is sensed by the system accelerometer. When this sensing occurs, the 0.05g light should come on, the scroll should begin to drive, and the lrange-to-go counter should begin to count down. The crew reported the light failure but were unable to verify whether or not the scroll or counter responded before the switch was manually changed to the backup mode. + +The failure had to be in the light, in the 0.05g sensing mechanism, or in the mode switch, mode switching could also have been premature. +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Either a short between the temperature switch wires to ground or a contamination in the auxiliary relay would actuate the light. The shorted condition could have resulted from electrolyte shorting within the battery case associated with the current surges discussed in the previous section. Contamination of the auxiliary relay has occurred in the past, and relays already packaged were not retrofitted since a false over-temperature indication can be identified as it was here. + +Corrective action is being taken to prevent electrolyte shorts associated with the previously discussed battery anomaly which should eliminate this type of sensor problem in future spacecraft. No further corrective action to eliminate contamination in the auxiliary relay is required. + +This anomaly is closed. + +14.2.4 Ascent Oxygen Tank 2 Shutoff Valve Leak +------ +Result 2: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ + +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +An investigation conducted after Apollo l2 did not identify a definite source of the problem, since extreme corrosion from sea water after landing prevented meaningful examination of the mechanical drive system and restricted testing to the power and servo assembly which contains the major electronic components. No abnormal indications were found in the Apollo l2 system; however, the failure symptoms were reproduced on a breadboard by breaking down the isolation across a transformer in the tachometer feedback compensation network. Although depotting and testing of the actual transformer failed to produce any evidence of malfunction, this mechanism was considered a likely candidate for a random failure. +------ +Result 2: +Performance of the guidance, navigation, and control system was normal except for two instances. Random motion observed in the sextant shaft during the zero optics mode was operationally prevented by turning off power to the optical system when not in use. This problem occurred during Apollo l2 and is thought to be caused by a buildup of contact resistance in the slip rings of the half-speed resolver in the sextant (section 14.l.3). The crew reported the 0.05g light did not illuminate as required within 3 seconds after the digital computer had indicated 0.05g. A manual procedure was therefore required to start the entry monitor system, which performed nominally throughout the remainder of entry (section 14.l.5). As a result of the aborted mission, all power was removed from the inertial platform, including heaters, for approximately 80 hours. After powering up and coarse aligning the platform to that of the lunar module, the cormmand module was guided to a successful landing within +------ + +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Performance of the guidance, navigation, and control system was normal except for two instances. Random motion observed in the sextant shaft during the zero optics mode was operationally prevented by turning off power to the optical system when not in use. This problem occurred during Apollo l2 and is thought to be caused by a buildup of contact resistance in the slip rings of the half-speed resolver in the sextant (section 14.l.3). The crew reported the 0.05g light did not illuminate as required within 3 seconds after the digital computer had indicated 0.05g. A manual procedure was therefore required to start the entry monitor system, which performed nominally throughout the remainder of entry (section 14.l.5). As a result of the aborted mission, all power was removed from the inertial platform, including heaters, for approximately 80 hours. After powering up and coarse aligning the platform to that of the lunar module, the cormmand module was guided to a successful landing within +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:47:02 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThis anomaly is open.\n\n14.1.5 Entry Monitor System 0.05g Light Malfunction\n\nT...', 'Result 1:\nEither a short between the temperature switch wires to ground or a contaminat...', 'Result 1:\nAn investigation conducted after Apollo l2 did not identify a definite source...', 'Result 1:\nPerformance of the guidance, navigation, and control system was normal except...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:47:02 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nstarted to sight the service module in the docking window. The lightened spac...', 'Result 1:\nAt the time of the oxygen tank incident, three events took place that affecte...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nPlatform-sensed velocity changes, ft/sec Command module axes Lunar module axe...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...'] +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ + +2025-04-03 at 19:47:02 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe two interrupters open the electrical circuits about 30 milliseconds befor...'] +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +2025-04-03 at 19:47:02 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:47:02 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.50 +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:47:02 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:47:02 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.140, Max reward: 0.642 +2025-04-03 at 19:47:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:47:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 lunar module power status report" +2025-04-03 at 19:47:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:47:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"lunar module battery capacity at lunar module undocking apollo mission" + +2025-04-03 at 19:47:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:47:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module batteries capacity ah upon undock +2025-04-03 at 19:47:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:47:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the amount of ampere hours (Ah) of power remaining in the Apollo Lunar Module's batteries at undocking? + + +2025-04-03 at 19:47:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:47:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission lunar module power consumption lunar surface return efficiency calculations +2025-04-03 at 19:47:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:47:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:47:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module battery capacity at undock +2025-04-03 at 19:47:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:47:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the estimated total ampere hours remaining in the lunar module's batteries at the time of undocking? + + +2025-04-03 at 19:47:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-03 at 19:47:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module ascent battery power consumption 12-ampere level +2025-04-03 at 19:47:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:47:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:47:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the exact difference in ampere hours between the total available and total used energy in the descent and ascent-stage batteries? + + +2025-04-03 at 19:47:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:47:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:47:17 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:47:17 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:47:17 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, True, False, False] +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_correctness:82 - Student lengths: [154, 380, 665, 226, 398, 531] +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [3, 3, 3, 3, 3, 3] +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_correctness:84 - Average student length: 392.33 +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 3.00 +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_correctness:86 - Length ratio: 130.78 +2025-04-03 at 19:47:17 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:47:17 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.371 Âą 0.198 +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.17 Âą 1.34 +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 2/6 +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [4, 0, 1, 2, 3, 3] +2025-04-03 at 19:47:17 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:47:17 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe electrical power system performed all required functions. At lunar module...'] +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 19:47:17 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:47:17 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...'] +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:47:17 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...'] +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + +Figure 7.2-l.- Lunar module water usage. + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Figure ll.l-2.- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:47:17 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...'] +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 1.0 +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.167 +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:47:17 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.367, Perfect scores: 0/6 +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.67 +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 2/6 +2025-04-03 at 19:47:17 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:47:17 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.359, Max reward: 0.808 +2025-04-03 at 19:47:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:47:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: acceleration drift in mERU/g for input axis in Z-direction +2025-04-03 at 19:47:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.īŋŊīŋŊ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:47:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: mERU/g acceleration drift electric vehicle motor Z-axis input axis +2025-04-03 at 19:47:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:47:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: mercedes ETG specifications acceleration z direction +2025-04-03 at 19:47:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 19:47:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'acceleration drift in z-axis material-specific effects mERU/g' +2025-04-03 at 19:47:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:47:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:47:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system Deep Space 1 accelerometer bias error acceleration drift mERU/g +2025-04-03 at 19:47:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-03 at 19:47:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: meru/g conversion to g acceleration units +2025-04-03 at 19:47:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:47:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'acceleration drift z-axis input axis measurement data comparisons' +2025-04-03 at 19:47:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:47:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:47:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system Deep Space 1 6.5 REACTION CONTROL accelerometer bias acceleration drift Z-axis +2025-04-03 at 19:47:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ + +2025-04-03 at 19:47:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: typical acceleration drift for acceleration in mERU/g for input axis z direction +2025-04-03 at 19:47:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:47:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'acceleration drift z-axis input axis commanded platform discrepancy ' +2025-04-03 at 19:47:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:47:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:47:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system Deep Space 1 yawing angle deviation mERU/g acceleration drift +2025-04-03 at 19:47:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ +Result 2: +The passive thermal control mode was reestablished by rolling 90 degrees with reference to the abort-guidance-driven attitude displays. This maneuver placed the terminator parallel to the X-axis of the crewmen optical alignment sight. Rates were nulled in pitch and roll with the thrust/ translation controller assembly. Yaw was again automatically controlled by the abort guidance system. Nulling rates to zero was impossible because of the inaccurate readout of the rate needles. When rates appeared to be nulled, yaw control was placed in the reaction control pulse mode. Twelve yaw-right pulses were then used to start the passive thermal control mode maneuver. Because rates could not be completely nulled, some roll-pitch coupling was observed. +------ + +2025-04-03 at 19:47:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:47:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system Deep Space 1 yawing mode yaw control mERU/g acceleration drift +2025-04-03 at 19:47:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The passive thermal control mode was reestablished by rolling 90 degrees with reference to the abort-guidance-driven attitude displays. This maneuver placed the terminator parallel to the X-axis of the crewmen optical alignment sight. Rates were nulled in pitch and roll with the thrust/ translation controller assembly. Yaw was again automatically controlled by the abort guidance system. Nulling rates to zero was impossible because of the inaccurate readout of the rate needles. When rates appeared to be nulled, yaw control was placed in the reaction control pulse mode. Twelve yaw-right pulses were then used to start the passive thermal control mode maneuver. Because rates could not be completely nulled, some roll-pitch coupling was observed. +------ +Result 2: +attitude error needles to maintain attitude. Attitude control during the maneuver was performed by manually nulling the pitch and roll error needles. This maneuver necessarily required crew-cooperation, since the Lunar Module Pilot controlled pitch and the Commander controlled roll. Yaw attitude was maintained automatically by the abort guidance system. The Command Module Pilot called out the engine start and stop times, and the entire l4-second firing was performed at l0 percent thrust. The engine was shut down l second short of the calculated firing time to preclude an overburn which might require use of minus-X thrusters and cause plume impingement on the command module. The control and alignment techniques to accomplish such a contingency midcourse maneuver are believed to be satisfactory. +------ + +2025-04-03 at 19:47:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:47:32 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:47:32 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:47:32 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, True, False, False] +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_correctness:82 - Student lengths: [562, 801, 513, 107, 112, 806] +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_correctness:84 - Average student length: 483.50 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 4.00 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_correctness:86 - Length ratio: 120.88 +2025-04-03 at 19:47:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:47:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.433 Âą 0.352 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.00 Âą 1.83 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [1, 0, 5, 3, 3, 0] +2025-04-03 at 19:47:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:47:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...'] +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +2025-04-03 at 19:47:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ + +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ +Result 2: +The passive thermal control mode was reestablished by rolling 90 degrees with reference to the abort-guidance-driven attitude displays. This maneuver placed the terminator parallel to the X-axis of the crewmen optical alignment sight. Rates were nulled in pitch and roll with the thrust/ translation controller assembly. Yaw was again automatically controlled by the abort guidance system. Nulling rates to zero was impossible because of the inaccurate readout of the rate needles. When rates appeared to be nulled, yaw control was placed in the reaction control pulse mode. Twelve yaw-right pulses were then used to start the passive thermal control mode maneuver. Because rates could not be completely nulled, some roll-pitch coupling was observed. +------ + +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The passive thermal control mode was reestablished by rolling 90 degrees with reference to the abort-guidance-driven attitude displays. This maneuver placed the terminator parallel to the X-axis of the crewmen optical alignment sight. Rates were nulled in pitch and roll with the thrust/ translation controller assembly. Yaw was again automatically controlled by the abort guidance system. Nulling rates to zero was impossible because of the inaccurate readout of the rate needles. When rates appeared to be nulled, yaw control was placed in the reaction control pulse mode. Twelve yaw-right pulses were then used to start the passive thermal control mode maneuver. Because rates could not be completely nulled, some roll-pitch coupling was observed. +------ +Result 2: +attitude error needles to maintain attitude. Attitude control during the maneuver was performed by manually nulling the pitch and roll error needles. This maneuver necessarily required crew-cooperation, since the Lunar Module Pilot controlled pitch and the Commander controlled roll. Yaw attitude was maintained automatically by the abort guidance system. The Command Module Pilot called out the engine start and stop times, and the entire l4-second firing was performed at l0 percent thrust. The engine was shut down l second short of the calculated firing time to preclude an overburn which might require use of minus-X thrusters and cause plume impingement on the command module. The control and alignment techniques to accomplish such a contingency midcourse maneuver are believed to be satisfactory. +------ + +2025-04-03 at 19:47:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nCondition Maneuver Second midcourse correction Transearth injection Third mid...', 'Result 1:\nUncompens ated Error term error One-sigma specification Offset velocity, ft/s...', 'Result 1:\nUncompens ated Error term error One-sigma specification Offset velocity, ft/s...', 'Result 1:\nTo conserve reaction control fuel when holding an attitude, a wide deadband w...', 'Result 1:\nThe passive thermal control mode was reestablished by rolling 90 degrees with...'] +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:47:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nParame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .6...', 'Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...', 'Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...'] +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:47:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...', 'Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...', 'Result 1:\nTo conserve reaction control fuel when holding an attitude, a wide deadband w...'] +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu" after ugdut: Accelerometera X-Scule factor error、ppm. 2 -19y 24 7 -199 BiaīŧŒcm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1īŧŸ +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axisīŧŒmERU/g..īŧŒ -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axiæ—Ĩ,mERU/g +19.17 7.14 7 +2025-04-03 at 19:47:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:47:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.367, Perfect scores: 1/6 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 4/6 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.17 +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 19:47:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:47:32 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.338, Max reward: 0.750 +2025-04-03 at 19:47:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:47:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch weather conditions +2025-04-03 at 19:47:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 19:47:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 launch weather conditions +2025-04-03 at 19:47:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:47:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: weather forecast kennedy space center july 11 1970 +2025-04-03 at 19:47:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 19:47:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:47:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: conditions at cape canaveral launch site during apollo 13 +2025-04-03 at 19:47:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:47:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 launch day lightning storm" or "Apollo 13 mission launch weather conditions" +2025-04-03 at 19:47:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-03 at 19:47:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: weather conditions april 11 1970 kennedy space center +2025-04-03 at 19:47:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:47:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:47:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: bad weather conditions cape canaveral during apollo 13 launch +2025-04-03 at 19:47:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:47:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollos 12 and 13 launch electrical disturbances" or "NASA research on launch phase electrical phenomena" +2025-04-03 at 19:47:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 19:47:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: april 11 1970 kennedy space center weather +2025-04-03 at 19:47:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:47:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:47:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar launch site electrical charge warnings +2025-04-03 at 19:47:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 19:47:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 launch electrical charge separation" or "Launch electrical hazards assessment Apollo 13" +2025-04-03 at 19:47:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-03 at 19:47:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: april 13 1970 kennedy space center weather +2025-04-03 at 19:47:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:47:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:47:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: thumping noise lunar module Apollo 13 +2025-04-03 at 19:47:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:47:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 launch electrostatic potential" or "electrostatic hazards Apollo 13 mission" +2025-04-03 at 19:47:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-03 at 19:47:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: april 1970 kennedy space center weather +2025-04-03 at 19:47:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:47:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:47:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: short circuit lunar module apollo 13 +2025-04-03 at 19:47:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:47:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: april 11 1970 weather houston +2025-04-03 at 19:47:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +159.70E 159.56E 188 371.38 188 393.19 3 065.8 3 093.2 79.364 79.934 115.464 116.54 Transearth phase Transearth injection Ignition Cutoff Thirdmidcourse correction Moon Moon 79:27 :39.0 79:32:02.8 3.73N 3.62N 65.46E 64.77E 5 465.26 5 658.68 4 547.7 5 020.2 72.645 64.784 -116.308 -117.886 Ignition Earth Earth 105:18:28.0 105:18:42.0 19.63N 19.50N 136.84W 136.90W 152 224.32 152 215.52 4 457.8 4456.6 -79.673 -79.765 114.134 114.242 Fourthmidcourse correction Ignition Cutofr Earth Earth 137:39:51.5 137:40:13.0 11.35N 11.34N 113.39E 113.32E 37 806.58 37 776.05 10 109.1 10 114.6 -72.369 -72.373 116.663 118.660 Service module separation Earth 138:01:48.0 10.88N 108.77E 35 694.93 10405.9 -71.941 118.824 Undocking Earth 141:30:00.2 1.23S 77.55E 11 257.48 1.7 465.9 -60.548 120.621 Entry interface Earth 142:40:45.7 28.23S 173.44E 65.83 36 210.6 -6.269 77.210 +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 19:47:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:47:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch weather conditions +2025-04-03 at 19:47:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 19:47:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: april 11 1970 launch hurricane +2025-04-03 at 19:47:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 19:47:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:48:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lightning storm +2025-04-03 at 19:48:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-03 at 19:48:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: april 1970 weather kennedy space center +2025-04-03 at 19:48:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:48:00 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:48:00 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:48:00 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1511, 301, 389, 524, 1246, 308] +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_correctness:84 - Average student length: 713.17 +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 8.00 +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_correctness:86 - Length ratio: 89.15 +2025-04-03 at 19:48:00 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_format:228 - Responses ending properly: 0/6 +2025-04-03 at 19:48:00 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.317 Âą 0.354 +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 5.67 Âą 7.23 +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 2/6 +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [20, 0, 5, 0, 9, 0] +2025-04-03 at 19:48:00 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down. + + + +Figure 14-l5.- Descent stage battery location. + + + +The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Equipment transfer and stowage in both the command module and lunar module was completed about 7 hours prior to entry, with the exception of the cameras that were to be used for service module photography. At 6-1/2 hours before entry, command module activity included powering up the instrumentation and placing entry battery C on main bus A, with main bus B still powered from the lunar module. The command module reaction control thrusters were preheated for 20 minutes, and all instrumented engines were observed to be above the minimum operating temperature l0 minutes after heater operation was terminated. + +8.9.3 Final Midcourse Correction +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +2025-04-03 at 19:48:00 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +Field meter records indicate the Apollo l3 vehicle carried aloft a net positive charge and that the trailing exhaust gases were negatively charged (fig. ll.l-4). Initial analysis indicates the total charge $\mathsf Q$ carried by the vehicle was about 0.o4 coulomb. If the capacitance of the launch vehicle is about l00 picofarads, the vehicle is then at a potential of 4 million volts. A stored charge of $0.04$ coulomb at a potential of 4 million volts provides an electrostatic potential energy of 160 0o0 joules. Although this energy is much less than that dissipated in a natural lightning discharge, the level is still considerable and could significantly increase the potential hazard in an otherwise marginal weather situation. These numbers are consistent with the electrostatic discharge analysis performed on the Apollo l2 lightning incident. Engines in jet aircraft have been observed to produce similar charging effects . + + + +Figure ll.l-4.- Electrical charge characteristics. +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +2025-04-03 at 19:48:00 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +159.70E 159.56E 188 371.38 188 393.19 3 065.8 3 093.2 79.364 79.934 115.464 116.54 Transearth phase Transearth injection Ignition Cutoff Thirdmidcourse correction Moon Moon 79:27 :39.0 79:32:02.8 3.73N 3.62N 65.46E 64.77E 5 465.26 5 658.68 4 547.7 5 020.2 72.645 64.784 -116.308 -117.886 Ignition Earth Earth 105:18:28.0 105:18:42.0 19.63N 19.50N 136.84W 136.90W 152 224.32 152 215.52 4 457.8 4456.6 -79.673 -79.765 114.134 114.242 Fourthmidcourse correction Ignition Cutofr Earth Earth 137:39:51.5 137:40:13.0 11.35N 11.34N 113.39E 113.32E 37 806.58 37 776.05 10 109.1 10 114.6 -72.369 -72.373 116.663 118.660 Service module separation Earth 138:01:48.0 10.88N 108.77E 35 694.93 10405.9 -71.941 118.824 Undocking Earth 141:30:00.2 1.23S 77.55E 11 257.48 1.7 465.9 -60.548 120.621 Entry interface Earth 142:40:45.7 28.23S 173.44E 65.83 36 210.6 -6.269 77.210 +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:48:00 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...', 'Result 1:\nThe space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, ...', 'Result 1:\nThe space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, ...', 'Result 1:\nThe space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, ...', 'Result 1:\nThe space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, ...', 'Result 1:\n159.70E 159.56E 188 371.38 188 393.19 3 065.8 3 093.2 79.364 79.934 115.464 1...', 'Result 1:\nThe space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, ...', 'Result 1:\nThe space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, ...'] +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +2025-04-03 at 19:48:00 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:48:00 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.267, Perfect scores: 0/6 +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.67 +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:48:00 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:48:00 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.180, Max reward: 0.487 +2025-04-03 at 19:48:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:48:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "vehicle launch azimuth +2025-04-03 at 19:48:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 19:48:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the azimuth of the Apollo 11 lunar module launch from Kennedy Space Center? +2025-04-03 at 19:48:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 19:48:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the launch azimuth of the Mars 2020 Perseverance Rover launch? +2025-04-03 at 19:48:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ + +2025-04-03 at 19:48:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vehicle launch direction or spacecraft launch azimuth +2025-04-03 at 19:48:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 19:48:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:48:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 14 launch azimuth +2025-04-03 at 19:48:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:48:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the launch azimuth and elevation angle at the S-IVB engine separation for Apollo 13? +2025-04-03 at 19:48:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 19:48:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the specific launch azimuth of the Perseverance rover's launch, specifically the one on July 20, 2020? +2025-04-03 at 19:48:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 19:48:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:48:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 launch azimuth July 1969 +2025-04-03 at 19:48:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:48:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 AS08-22-2164 +2025-04-03 at 19:48:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:48:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the launch azimuth of the Perseverance rover's launch, specifically on July 20, 2020? +2025-04-03 at 19:48:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 19:48:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:48:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 launch trajectory +2025-04-03 at 19:48:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:48:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch azimuth and roll maneuver time +2025-04-03 at 19:48:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:48:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the launch azimuth of the Perseverance rover's launch, specifically on July 20, 2020? +2025-04-03 at 19:48:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 19:48:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:48:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 launch azimuth results +2025-04-03 at 19:48:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:48:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission launch timeline and azimuth +2025-04-03 at 19:48:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:48:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the launch date and azimuth of the Perseverance rover's launch? +2025-04-03 at 19:48:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 19:48:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:48:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 azimuth of rocket launch +2025-04-03 at 19:48:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-03 at 19:48:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 launch azimuth and roll maneuver time May 1970 +2025-04-03 at 19:48:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:48:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the launch date and azimuth of the Perseverance rover's launch? +2025-04-03 at 19:48:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 19:48:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:48:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 lunar module launch azimuth +2025-04-03 at 19:48:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-03 at 19:48:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What to know about Perseverance rover launch? +2025-04-03 at 19:48:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 19:48:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:48:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 launch azimuth May 1969 +2025-04-03 at 19:48:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:48:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the launch date and azimuth of the Perseverance rover's launch? +2025-04-03 at 19:48:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 19:48:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:48:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 S-IVB launch azimuth April 1969 +2025-04-03 at 19:48:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:48:20 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:48:20 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:48:20 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, False, False] +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_correctness:82 - Student lengths: [196, 1898, 2011, 49, 184, 325] +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [24, 24, 24, 24, 24, 24] +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_correctness:84 - Average student length: 777.17 +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 24.00 +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_correctness:86 - Length ratio: 32.38 +2025-04-03 at 19:48:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:48:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.533 Âą 0.433 +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 4.00 Âą 3.79 +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [6, 9, 8, 0, 1, 0] +2025-04-03 at 19:48:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 19:48:20 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 19:48:20 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:48:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.17 +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:48:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:48:20 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.209, Max reward: 0.708 +2025-04-03 at 19:48:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:48:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "when and where were the first balloons launched +2025-04-03 at 19:48:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 19:48:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What were the first balloons to be launched into space and their launch time? +2025-04-03 at 19:48:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:48:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "when were the first two balloons launched" +2025-04-03 at 19:48:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 19:48:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "first high-altitude balloons launch date time" +2025-04-03 at 19:48:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 19:48:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:48:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo 13 time of launch of balloons" and "when did the first two balloons for the aeOS mission launch" +2025-04-03 at 19:48:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:48:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: " launch time of three weather balloons 1970 April " +2025-04-03 at 19:48:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 19:48:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "High-altitude balloons first launch times" +2025-04-03 at 19:48:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 19:48:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:48:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "first two balloons launched with Apollos program" +2025-04-03 at 19:48:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:48:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo 11 launch time balloons experiment" +2025-04-03 at 19:48:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +The Apollo l3 space vehicle was launched from pad A of complex 39, Kennedy Space Center, Florida. Except for the high-amplitude, lowfrequency oscillations which resulted in premature cutoff of the S-II center engine, the basic performance of the launch vehicle was normal. Despite the anomaly, all launch vehicle objectives were achieved, as discussed in reference 2. In addition, the S-IvB lunar impact experiment was accomplished, as discussed in section ll.3. +------ + +2025-04-03 at 19:48:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:48:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "first balloons launched apr 1970" +2025-04-03 at 19:48:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:48:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo 11 launch time relative to pilot describing function experiment" +2025-04-03 at 19:48:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +The Apollo l3 space vehicle was launched from pad A of complex 39, Kennedy Space Center, Florida. Except for the high-amplitude, lowfrequency oscillations which resulted in premature cutoff of the S-II center engine, the basic performance of the launch vehicle was normal. Despite the anomaly, all launch vehicle objectives were achieved, as discussed in reference 2. In addition, the S-IvB lunar impact experiment was accomplished, as discussed in section ll.3. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:48:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:48:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo mission 11 launch time three weather balloons experiment" +2025-04-03 at 19:48:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +The Apollo l3 space vehicle was launched from pad A of complex 39, Kennedy Space Center, Florida. Except for the high-amplitude, lowfrequency oscillations which resulted in premature cutoff of the S-II center engine, the basic performance of the launch vehicle was normal. Despite the anomaly, all launch vehicle objectives were achieved, as discussed in reference 2. In addition, the S-IvB lunar impact experiment was accomplished, as discussed in section ll.3. +------ + +2025-04-03 at 19:48:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:48:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo 13 launch time three weather balloons" +2025-04-03 at 19:48:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:48:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:48:37 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:48:37 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:48:37 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_correctness:82 - Student lengths: [288, 55, 615, 435, 780, 136] +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [13, 13, 13, 13, 13, 13] +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_correctness:84 - Average student length: 384.83 +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 13.00 +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_correctness:86 - Length ratio: 29.60 +2025-04-03 at 19:48:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:48:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.433 Âą 0.362 +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.17 Âą 2.19 +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [4, 1, 6, 0, 0, 2] +2025-04-03 at 19:48:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +The Apollo l3 space vehicle was launched from pad A of complex 39, Kennedy Space Center, Florida. Except for the high-amplitude, lowfrequency oscillations which resulted in premature cutoff of the S-II center engine, the basic performance of the launch vehicle was normal. Despite the anomaly, all launch vehicle objectives were achieved, as discussed in reference 2. In addition, the S-IvB lunar impact experiment was accomplished, as discussed in section ll.3. +------ + +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +The Apollo l3 space vehicle was launched from pad A of complex 39, Kennedy Space Center, Florida. Except for the high-amplitude, lowfrequency oscillations which resulted in premature cutoff of the S-II center engine, the basic performance of the launch vehicle was normal. Despite the anomaly, all launch vehicle objectives were achieved, as discussed in reference 2. In addition, the S-IvB lunar impact experiment was accomplished, as discussed in section ll.3. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +The Apollo l3 space vehicle was launched from pad A of complex 39, Kennedy Space Center, Florida. Except for the high-amplitude, lowfrequency oscillations which resulted in premature cutoff of the S-II center engine, the basic performance of the launch vehicle was normal. Despite the anomaly, all launch vehicle objectives were achieved, as discussed in reference 2. In addition, the S-IvB lunar impact experiment was accomplished, as discussed in section ll.3. +------ + +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-03 at 19:48:37 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-03 at 19:48:37 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the "fair weather" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an + +altitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood. + +11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:48:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.033, Perfect scores: 0/6 +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:48:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:48:37 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.209, Max reward: 0.743 +2025-04-03 at 19:48:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:48:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: WGS84 Y-axis error magnitude comparison regional mapping missions standard geodetic observations +2025-04-03 at 19:48:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:48:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Y-axis error magnitude discrepancy in measurement instruments" +2025-04-03 at 19:48:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:48:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:48:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: magnitude of null bias drift during launch phase +2025-04-03 at 19:48:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:48:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:48:51 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:48:51 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:48:51 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, False, False, False] +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_correctness:82 - Student lengths: [411, 648, 918, 879, 613, 564] +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [36, 36, 36, 36, 36, 36] +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_correctness:84 - Average student length: 672.17 +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 36.00 +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_correctness:86 - Length ratio: 18.67 +2025-04-03 at 19:48:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:48:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.137 Âą 0.198 +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.17 Âą 2.19 +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 0, 6, 0, 0, 1] +2025-04-03 at 19:48:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-03 at 19:48:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-03 at 19:48:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-03 at 19:48:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-03 at 19:48:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:48:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:48:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:48:51 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.115, Max reward: 0.487 +2025-04-03 at 19:48:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:48:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "oldest recorded data point in history or ancient times" +2025-04-03 at 19:48:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 19:48:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the earliest recorded data point in astronomy or medical history? +2025-04-03 at 19:48:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 19:48:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "first recorded data point time range" +2025-04-03 at 19:48:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:48:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Oldest recorded temperature" + +2025-04-03 at 19:48:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 19:48:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:49:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Early mission log entry or data point Apollo 11. +2025-04-03 at 19:49:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:49:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-IVB impact seismic records +2025-04-03 at 19:49:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:49:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"lunar surface instrumentation first recorded data point" + +2025-04-03 at 19:49:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:49:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:49:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 transearth injection earliest recorded log entry. +2025-04-03 at 19:49:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 19:49:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-IVB impact time +2025-04-03 at 19:49:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 19:49:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"first moonquake seismic wave duration" + +2025-04-03 at 19:49:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 19:49:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:49:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Transearth injection, Apollo 11 mission, exact time. +2025-04-03 at 19:49:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:49:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:49:07 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:49:07 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:49:07 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, True, False, True] +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_correctness:82 - Student lengths: [855, 439, 575, 319, 1009, 510] +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_correctness:84 - Average student length: 617.83 +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 5.00 +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_correctness:86 - Length ratio: 123.57 +2025-04-03 at 19:49:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 19:49:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.408 Âą 0.318 +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.83 Âą 1.57 +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [1, 0, 4, 3, 0, 3] +2025-04-03 at 19:49:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ +Result 2: +Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off. + + + +Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch. + + + +Figure 1l.l-3.- Concluded + +No significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations. +------ + +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 19:49:07 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as "pinpoint novas ," "roman candles," and "similar to traces in a cloud chamber." More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute. + +9.3 PHYSICAL EXAMINATIONS + +Preflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary. +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:49:07 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe crew reported seeing point flashes or streaks of light, as had been previ...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...'] +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 19:49:07 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 19:49:07 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...', 'Result 1:\nDescription Completed B C Television coverage No Contingency sample collectio...', 'Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...'] +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:49:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.033, Perfect scores: 0/6 +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:49:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:49:07 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.311, Max reward: 0.856 +2025-04-03 at 19:49:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:49:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Failure of service propulsion auxiliary propellant gaging system on a mission performance impact +2025-04-03 at 19:49:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:49:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: probe system failure propellant ganging system mission performance +2025-04-03 at 19:49:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:49:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:49:12 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:49:12 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:49:13 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, False, False, False, False] +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_correctness:82 - Student lengths: [377, 618, 463, 471, 429, 146] +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [2, 2, 2, 2, 2, 2] +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_correctness:84 - Average student length: 417.33 +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 2.00 +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_correctness:86 - Length ratio: 208.67 +2025-04-03 at 19:49:13 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:49:13 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.117 Âą 0.165 +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.33 Âą 0.47 +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 1, 0, 1, 0, 0] +2025-04-03 at 19:49:13 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-03 at 19:49:13 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-03 at 19:49:13 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-03 at 19:49:13 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +2025-04-03 at 19:49:13 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:49:13 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.033, Perfect scores: 0/6 +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:49:13 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:49:13 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.067, Max reward: 0.200 +2025-04-03 at 19:49:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:49:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"pericynthion altitude translunar injection Apollo 11" +2025-04-03 at 19:49:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pericynthion altitude translunar injection +2025-04-03 at 19:49:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 19:49:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Perihelion altitude at translunar injection +2025-04-03 at 19:49:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 19:49:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the achieved pericynthion altitude at translunar injection for Apollo 11 mission? +2025-04-03 at 19:49:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:49:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:49:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 translunar injection pericynthion altitude +2025-04-03 at 19:49:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:49:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 free-return pericynthion altitude +2025-04-03 at 19:49:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:49:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 free-return pericynthion altitude at end of maneuver +2025-04-03 at 19:49:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:49:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 free-return pericynthion altitude at 2 hours +2025-04-03 at 19:49:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:49:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 free-return pericynthion altitude after 2 hours +2025-04-03 at 19:49:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:49:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 free-return pericynthion altitude at 2+ hours +2025-04-03 at 19:49:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:49:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 free-return pericynthion altitude +2025-04-03 at 19:49:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:49:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 pericynthion altitude 2 hours after translunar injection +2025-04-03 at 19:49:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:31 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:49:31 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:49:31 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, False, True, True, False] +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_correctness:82 - Student lengths: [444, 599, 1631, 319, 234, 225] +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [11, 11, 11, 11, 11, 11] +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_correctness:84 - Average student length: 575.33 +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 11.00 +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_correctness:86 - Length ratio: 52.30 +2025-04-03 at 19:49:31 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 19:49:31 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.333 Âą 0.317 +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.00 Âą 3.16 +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 1, 9, 1, 1, 0] +2025-04-03 at 19:49:31 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 19:49:31 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +2025-04-03 at 19:49:31 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:49:31 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 19:49:31 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:49:31 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.100, Max reward: 0.200 +2025-04-03 at 19:49:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:49:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command and service module rotation moment of inertia +2025-04-03 at 19:49:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:49:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the moment of inertia of the command service module +2025-04-03 at 19:49:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:49:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command and service module moment of inertia +2025-04-03 at 19:49:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 19:49:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Technical Reports Server "ISS Service Module" OR "ISS Command Module" moment of inertia +2025-04-03 at 19:49:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 19:49:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:49:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: definition angular moment of inertia of command module +2025-04-03 at 19:49:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 19:49:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moment of inertia of the International Space Station Service Module +2025-04-03 at 19:49:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:49:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moment of inertia command service module spacecraft +2025-04-03 at 19:49:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 19:49:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "International Space Station moment of inertia Service Module" OR "ISS Command Module moment of inertia" +2025-04-03 at 19:49:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 19:49:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:49:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: translational inertia lunar module command module +2025-04-03 at 19:49:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:49:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moment of inertia of a spacecraft module definition +2025-04-03 at 19:49:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:49:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moment of inertia command service module apollo csm +2025-04-03 at 19:49:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 19:49:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "moment of inertia of International Space Station Command Module" OR "ISS Service Module mass moment of inertia" +2025-04-03 at 19:49:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:49:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:49:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: translational inertia command module lunar module +2025-04-03 at 19:49:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 19:49:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moment of inertia formula for a rectangular object +2025-04-03 at 19:49:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 19:49:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module docking velocity changes command service module +2025-04-03 at 19:49:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-03 at 19:49:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "SP-100 SSM inertial properties" OR "International Space Station Command Module mass properties" +2025-04-03 at 19:49:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:49:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:49:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: gravitational properties mass of command module +2025-04-03 at 19:49:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 19:49:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: mass and moment of inertia for a cylindrical object +2025-04-03 at 19:49:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-03 at 19:49:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command service module momentum inertia +2025-04-03 at 19:49:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:49:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "moment of inertia command and service module" OR "inertial properties lunar module" OR "Apollo module geometry and mass" +2025-04-03 at 19:49:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-03 at 19:49:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:49:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: mass of command module Apollo spacecraft +2025-04-03 at 19:49:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 19:49:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moment of inertia of the Intentional Space Station Command Service Module +2025-04-03 at 19:49:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 19:49:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module command service module inertial properties +2025-04-03 at 19:49:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 19:49:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:49:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moment of inertia calculated mass command module +2025-04-03 at 19:49:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:49:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: zero moment of inertia example +2025-04-03 at 19:49:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 19:49:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar service module inertia Apollo +2025-04-03 at 19:49:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:49:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:49:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moment of inertia calculation for body of revolution constant radius mass of command module +2025-04-03 at 19:49:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 19:49:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar command module inertia Apollo +2025-04-03 at 19:49:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 19:49:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:49:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module inertial constraints +2025-04-03 at 19:49:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:49:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:49:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar command module inertial drift +2025-04-03 at 19:49:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:49:58 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:49:58 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:49:58 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, False, False, False, False] +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_correctness:82 - Student lengths: [336, 1028, 308, 1887, 1829, 470] +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_correctness:84 - Average student length: 976.33 +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 5.00 +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_correctness:86 - Length ratio: 195.27 +2025-04-03 at 19:49:58 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 19:49:58 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.633 Âą 0.448 +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 5.00 Âą 3.83 +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 8, 0, 7, 10, 5] +2025-04-03 at 19:49:58 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 19:49:58 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 19:49:58 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Earth-centered inertial coordinates. Bystem. **Change in velocity showm in body X-axis for descent propulsion firings under control of abort guidance + +The primary guidance system was powered up at 133-l/2 hours, after which a coarse alignment to the abort guidance system was performed. The spacecraft axes had previously been aligned to an inertial reference using the abort guidance system by sighting on the earth with the crew optical alignment sight. Alignment accuracy was refined by performing a realignment using the sun and moon as sighting targets for the alignment optical telescope. The star-angle difference of minus l.l2 degrees resulted almost entirely from approximations in stored lunar and solar ephemeris data and computer routines used to calculate sun and moon position vectors. + +6.4.4 Inertial Measurement Unit +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +A.1 COMMAND AND SERVICE MODULES +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:49:58 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nA.1 COMMAND AND SERVICE MODULES\n------\nResult 2:\nstarted to sight the service...', 'Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nstarted to sight the service module in the docking window. The lightened spac...', 'Result 1:\nDESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNA...', 'Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...', 'Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...', 'Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...', 'Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...'] +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +Spacecraft mass properties for the Apollo l2 mission are summarized in table A-I. These data represent the conditions as determined from postflight analyses of expendable loadings and usage during the flight. Variations in spacecraft mass properties are determined for each significant mission phase from lift-off through landing. Expendables usage is based on reported real-time and postflight data as presented in other sections of this report. The weights and centers of gravity of the individual command and service modules and of the lunar module ascent and descent stages were measured prior to flight, and the inertia values were calculated. All changes incorporated after the actual weighing were monitored, and the spacecraft mass properties were updated. + +TABLE A-I.- MASS PROPERTIES +------ + +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 19:49:58 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:49:58 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:49:58 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.298, Max reward: 0.606 +2025-04-03 at 19:50:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:50:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module pitch adjustment Apollo 10 pitch-angle +2025-04-03 at 19:50:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:50:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the pitch angle for commanding module pilot antenna adjustment during apollo 13 +2025-04-03 at 19:50:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 19:50:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 antenna angle +2025-04-03 at 19:50:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:50:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:50:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 10 lunar module pitch-angle adjustment +2025-04-03 at 19:50:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:50:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module pitch angle during apollo 13 high-gain antenna acquisition adjustment +2025-04-03 at 19:50:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 19:50:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:50:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 10 mission pitch-angle adjustment +2025-04-03 at 19:50:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:50:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:50:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: john young command module apollo 10 pitch angle adjustement +2025-04-03 at 19:50:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:50:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:50:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: jack swigert command module apollo 10 pitch angle adjustement +2025-04-03 at 19:50:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +With primary guidance available, guidance system steering was manually followed to trim the spacecraft attitudes for the maneuver. Although the displayed attitudes looked favorable in comparison to groundsupplied and out-the-window readings , the primary guidance steering needles read full scale left in roll and yaw (section 6.4). At about 137 hours 40 minutes, the lunar module reaction control system was used to provide a 2.9-ft/sec velocity correction. The maneuver was completed using manual pitch and roll control and abort guidance yaw control in a manner similar to that for the previous midcourse correction. + +8.9.4 Service Module Separation and Photography +------ + +2025-04-03 at 19:50:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:50:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 10 cm pilot pitch angle adjustment +2025-04-03 at 19:50:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 19:50:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:50:12 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:50:12 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:50:12 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, True, False, False, False] +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_correctness:82 - Student lengths: [334, 404, 315, 196, 117, 187] +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_correctness:84 - Average student length: 258.83 +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_correctness:86 - Length ratio: 25.88 +2025-04-03 at 19:50:12 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:50:12 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.300 Âą 0.350 +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.50 Âą 2.14 +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [6, 0, 2, 0, 1, 0] +2025-04-03 at 19:50:12 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +With primary guidance available, guidance system steering was manually followed to trim the spacecraft attitudes for the maneuver. Although the displayed attitudes looked favorable in comparison to groundsupplied and out-the-window readings , the primary guidance steering needles read full scale left in roll and yaw (section 6.4). At about 137 hours 40 minutes, the lunar module reaction control system was used to provide a 2.9-ft/sec velocity correction. The maneuver was completed using manual pitch and roll control and abort guidance yaw control in a manner similar to that for the previous midcourse correction. + +8.9.4 Service Module Separation and Photography +------ + +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 19:50:12 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...'] +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 19:50:12 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 19:50:12 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:50:12 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...'] +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +2025-04-03 at 19:50:12 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 1.0 +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.167 +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:50:12 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.167, Perfect scores: 0/6 +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.00 +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:50:12 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:50:12 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.251, Max reward: 0.752 +2025-04-03 at 19:50:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:50:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 emergency abort +2025-04-03 at 19:50:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:50:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 emergency abort +2025-04-03 at 19:50:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:50:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 emergency abort +2025-04-03 at 19:50:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:50:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apoll mission emergency abort Apollo 13 +2025-04-03 at 19:50:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 19:50:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo mission emergency abort" +2025-04-03 at 19:50:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 19:50:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:50:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 oxygen tank explosion cause +2025-04-03 at 19:50:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 19:50:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 aborted mission navigation +2025-04-03 at 19:50:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:50:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 emergency abort rescue +2025-04-03 at 19:50:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:50:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ( Apollo 13 emergency abort procedures Descent Propulsion System midcourse correction) +2025-04-03 at 19:50:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:50:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:50:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 accident causes and timeline +2025-04-03 at 19:50:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:50:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 descent propulsion system oxygen tank 2 anomaly report +2025-04-03 at 19:50:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:50:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (Apollo 13 mission landing requirements power management} +2025-04-03 at 19:50:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 19:50:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:50:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 crew training geology field trip +2025-04-03 at 19:50:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 19:50:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 mission abort procedure +2025-04-03 at 19:50:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:50:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (Apollo mission with emergency abort lunar module descent propulsion system power management) +2025-04-03 at 19:50:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:50:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:50:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar geology exercises +2025-04-03 at 19:50:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 19:50:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (Apollo mission with emergencyabort problem makeshift power management) +2025-04-03 at 19:50:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:50:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:50:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar field geology experiments +2025-04-03 at 19:50:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-03 at 19:50:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (Apollo 12 vs Apollo 13 comparison mission airflow difference) +2025-04-03 at 19:50:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:50:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:50:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (Apollo 13 mission details recommended solutions after hydrogen tank problem) +2025-04-03 at 19:50:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:50:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:50:29 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:50:29 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:50:29 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, True, False, False] +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_correctness:82 - Student lengths: [325, 795, 346, 387, 103, 180] +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [9, 9, 9, 9, 9, 9] +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_correctness:84 - Average student length: 356.00 +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 9.00 +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_correctness:86 - Length ratio: 39.56 +2025-04-03 at 19:50:29 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:50:29 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.592 Âą 0.346 +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 3.33 Âą 2.56 +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [6, 0, 4, 2, 7, 1] +2025-04-03 at 19:50:29 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 19:50:29 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 5.0 +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.833 +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:50:29 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.267, Perfect scores: 0/6 +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 4/6 +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.33 +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:50:29 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:50:29 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.377, Max reward: 0.613 +2025-04-03 at 19:50:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:50:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: power source energy use apple iphone 12 +2025-04-03 at 19:50:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 19:50:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "energy remaining in original three entry batteries" no, but instead I will use "energy capacity lost in original 3 iPhone lithium-ion batteries" +2025-04-03 at 19:50:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ + +2025-04-03 at 19:50:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the amount of energy remaining in the entry batteries of the Apollo 13 spacecraft after the accident? +2025-04-03 at 19:50:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 19:50:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 lunar module battery power backup" +2025-04-03 at 19:50:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 19:50:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:50:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 command module entry battery type +2025-04-03 at 19:50:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 19:50:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo space suits battery voltage and power" +2025-04-03 at 19:50:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 19:50:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission entry batteries total power remaining +2025-04-03 at 19:50:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 19:50:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:50:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: l78 ampere-hours command module entry battery purpose +2025-04-03 at 19:50:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 19:50:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Puwt battery total power remaining Apollo 13 +2025-04-03 at 19:50:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:50:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:50:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module entry battery energy capacity +2025-04-03 at 19:50:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 19:50:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar module total power delivered during Apollo 13 +2025-04-03 at 19:50:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:50:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:50:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar module total energy delivered during Apollo 13 +2025-04-03 at 19:50:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:50:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:50:46 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:50:46 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:50:46 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, False, False] +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_correctness:82 - Student lengths: [32, 412, 139, 400, 788, 268] +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [15, 15, 15, 15, 15, 15] +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_correctness:84 - Average student length: 339.83 +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 15.00 +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_correctness:86 - Length ratio: 22.66 +2025-04-03 at 19:50:46 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:50:46 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.417 Âą 0.368 +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.17 Âą 1.86 +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [4, 0, 2, 0, 5, 2] +2025-04-03 at 19:50:46 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 19:50:46 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nDuring the mission, the fuel cells supplied approximately l20 kW-h of energy ...', 'Result 1:\nCommand module battery performance was acceptable throughout the mission. Ent...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...'] +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-03 at 19:50:46 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ + +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 19:50:46 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...'] +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-03 at 19:50:46 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ + +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:50:46 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...'] +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 19:50:46 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe electrical power system performed all required functions. At lunar module...'] +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:50:46 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.67 +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:50:46 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:50:46 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.415, Max reward: 1.000 +2025-04-03 at 19:50:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:50:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How many channels of biomedical monitoring systems are typical for a lunar lander? +2025-04-03 at 19:50:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 19:50:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How many biomedical signals monitored by Apollo Lunar Module +2025-04-03 at 19:50:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-03 at 19:50:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Biomedical monitoring systems lunar module apollo" + +2025-04-03 at 19:50:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 19:50:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +Biomedical monitoring spacecraft systems capacity Lunar Module telemetry system signals monitored + +2025-04-03 at 19:50:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 19:50:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:50:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"lunar module biomedical monitoring capabilities apollo" + +2025-04-03 at 19:50:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:50:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:50:54 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:50:54 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:50:54 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, True, True, False] +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_correctness:82 - Student lengths: [377, 325, 816, 256, 331, 427] +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [3, 3, 3, 3, 3, 3] +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_correctness:84 - Average student length: 422.00 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 3.00 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_correctness:86 - Length ratio: 140.67 +2025-04-03 at 19:50:54 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:50:54 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.258 Âą 0.190 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.83 Âą 0.69 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 1, 0, 1, 2, 1] +2025-04-03 at 19:50:54 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 19:50:54 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 19:50:54 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:50:54 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.300, Perfect scores: 0/6 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 4/6 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.17 +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 2/6 +2025-04-03 at 19:50:54 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:50:54 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.152, Max reward: 0.310 +2025-04-03 at 19:50:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:50:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "cryogenic oxygen transportation incident during combat or industrial operations" +2025-04-03 at 19:50:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 19:50:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: o "oxygen tank failures leading to mission aborts in space exploration" +2025-04-03 at 19:50:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 19:50:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happened to the primary mission after cryogenic oxygen tank malfunction +2025-04-03 at 19:50:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ +Result 2: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ + +2025-04-03 at 19:50:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "A-12 hypoxia warning divert decision" +2025-04-03 at 19:50:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The major medical concern, recogmized immediately after the abort decision, was the possibility of carbon dioxide buildup in the lunar module atmosphere. Since the physiological effects of increased carbon dioxide concentration are well known and readily recognizable with proper biomedical monitoring, the allowable limit of carbon dioxide buildup was increased from the nominal 7.6 to 15mm Hg. The carbon dioxide level was above 7.6mm Hg for only a 4-hour period, and no adverse physiological effects or degradation in crew performance resulted from this elevated concentration. Modified use of the lithium hydroxide cartridges (section 6.7) maintained the carbon dioxide partial pressure well below lmm Hg for the remainder of the flight. + +9.2.3 Sleep +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 19:50:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:51:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Need to find information about the Apollo mission and its alternative plans after the cryogenic oxygen tank incident, prior to aborting the primary mission +2025-04-03 at 19:51:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 19:51:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:51:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Need to find information about a specific incident involving cryogenic oxygen tanks in a military setting and the primary mission decision that followed +2025-04-03 at 19:51:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 19:51:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:51:04 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:51:04 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:51:04 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, False, False] +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_correctness:82 - Student lengths: [410, 757, 109, 951, 396, 329] +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [84, 84, 84, 84, 84, 84] +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_correctness:84 - Average student length: 492.00 +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 84.00 +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_correctness:86 - Length ratio: 5.86 +2025-04-03 at 19:51:04 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:51:04 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.267 Âą 0.225 +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.17 Âą 1.07 +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [3, 1, 1, 2, 0, 0] +2025-04-03 at 19:51:04 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ +Result 2: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ + +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The major medical concern, recogmized immediately after the abort decision, was the possibility of carbon dioxide buildup in the lunar module atmosphere. Since the physiological effects of increased carbon dioxide concentration are well known and readily recognizable with proper biomedical monitoring, the allowable limit of carbon dioxide buildup was increased from the nominal 7.6 to 15mm Hg. The carbon dioxide level was above 7.6mm Hg for only a 4-hour period, and no adverse physiological effects or degradation in crew performance resulted from this elevated concentration. Modified use of the lithium hydroxide cartridges (section 6.7) maintained the carbon dioxide partial pressure well below lmm Hg for the remainder of the flight. + +9.2.3 Sleep +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 19:51:04 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe major medical concern, recogmized immediately after the abort decision, w...'] +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-03 at 19:51:04 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +2025-04-03 at 19:51:04 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 19:51:04 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.167, Perfect scores: 1/6 +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.33 +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 19:51:04 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:51:04 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.250, Max reward: 0.789 +2025-04-03 at 19:51:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:51:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system +2025-04-03 at 19:51:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 19:51:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the specific function of the lunar module reaction control system's MOID system during lunar descent? +2025-04-03 at 19:51:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:51:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar Module Reaction Control System mission control purpose +2025-04-03 at 19:51:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:51:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo lunar module reaction control system" +2025-04-03 at 19:51:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:51:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"lunar module reaction control system NASA mission" +2025-04-03 at 19:51:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:51:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:51:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module reaction control system +2025-04-03 at 19:51:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 19:51:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the primary function of the lunar module reaction control system during lunar descent? +2025-04-03 at 19:51:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:51:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar Module Reaction Control System primary function +2025-04-03 at 19:51:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:51:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar Module Reaction Control System design" +2025-04-03 at 19:51:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:51:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Lunar Module Ascent Descent Propulsion System NASA 'apex'" +2025-04-03 at 19:51:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:51:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:51:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module fuel usage lunar landing +2025-04-03 at 19:51:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:51:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Lunar Module Descent Propulsion System established parameters +2025-04-03 at 19:51:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:51:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar Module Reaction Control System NASA description +2025-04-03 at 19:51:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 19:51:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar Module Reaction Control System technical specifications" +2025-04-03 at 19:51:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:51:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Lunar Module Reaction Control System primary function" +2025-04-03 at 19:51:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:51:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:51:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module power up command module power down +2025-04-03 at 19:51:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:51:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Lunar Module Descent Engine parameters +2025-04-03 at 19:51:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:51:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar Module Reaction Control System NASA definition +2025-04-03 at 19:51:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:51:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Lunar Module Reaction Control System NASA Apollo" +2025-04-03 at 19:51:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:51:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:51:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module power system failures +2025-04-03 at 19:51:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 19:51:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How did the lunar module use its descent engine to execute a free-return circumlunar trajectory? +2025-04-03 at 19:51:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ + +2025-04-03 at 19:51:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Lunar Module Reaction Control System description +2025-04-03 at 19:51:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 19:51:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Lunar Module control systems NASA Apollo mission" +2025-04-03 at 19:51:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:51:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:51:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module mechanical issues +2025-04-03 at 19:51:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 19:51:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How does the lunar module's descent engine enable a transearth injection maneuver? +2025-04-03 at 19:51:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:51:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Function of Lunar Module Reaction Control System +2025-04-03 at 19:51:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 19:51:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Lunar Module attitude control system" +2025-04-03 at 19:51:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:51:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:51:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module undocking command module +2025-04-03 at 19:51:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 19:51:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the effect of the transearth injection maneuver on the lunar module's orbital parameters? +2025-04-03 at 19:51:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:51:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:51:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module undocking difficulties +2025-04-03 at 19:51:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 19:51:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the significance of the Transearth injection maneuver in terms of navigation accuracy during the aborted Apollo 11 mission? +2025-04-03 at 19:51:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-03 at 19:51:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:51:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module undocking procedure +2025-04-03 at 19:51:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:51:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the relationship between the transearth injection maneuver and the entry flight-path angle? +2025-04-03 at 19:51:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:51:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:51:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module undocking entry procedures +2025-04-03 at 19:51:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:51:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the control mechanism used for the pitch and roll during the first transearth midcourse correction? +2025-04-03 at 19:51:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 19:51:29 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:51:29 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:51:29 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, True, False, True] +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1730, 2037, 552, 689, 445, 760] +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [28, 28, 28, 28, 28, 28] +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_correctness:84 - Average student length: 1035.50 +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 28.00 +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_correctness:86 - Length ratio: 36.98 +2025-04-03 at 19:51:29 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-03 at 19:51:29 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.742 Âą 0.349 +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 5.83 Âą 3.58 +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [10, 10, 0, 6, 3, 6] +2025-04-03 at 19:51:29 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:51:29 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...', 'Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...', 'Result 1:\nThe unprecedented powered-down state of the command module required generatio...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nAll mechanical systems functioned properly. One mechanical anomaly, however, ...', 'Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...', 'Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...', 'Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...', 'Result 1:\nA major flight control function, in adaition to the monitoring of systems sta...'] +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 19:51:29 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nThe transearth injection maneuver was performed on time, and the transearth c...', 'Result 1:\nThe transearth injection maneuver was performed on time, and the transearth c...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nCondition Maneuver Second midcourse correction Transearth injection Third mid...'] +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 19:51:29 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 19:51:29 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...'] +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:51:29 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...'] +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:51:29 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...'] +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:51:29 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.17 +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:51:29 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:51:29 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.209, Max reward: 0.707 +2025-04-03 at 19:51:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:51:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Flyers flown from Pago Pago after flight" +2025-04-03 at 19:51:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +------ + +2025-04-03 at 19:51:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "flight crew after Pago Pago, Samoa destination +2025-04-03 at 19:51:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 19:51:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"flight crew Pago Pago Samoa destination +2025-04-03 at 19:51:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 19:51:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: flights from Pago Pago, Samoa +2025-04-03 at 19:51:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 19:51:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "flight crew Pago Pago Samoa destination +2025-04-03 at 19:51:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 19:51:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:51:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 8 flight crew after Pago Pago, Samoa next destination +2025-04-03 at 19:51:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:51:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: summarize Iwo Jima recovery operations +2025-04-03 at 19:51:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 19:51:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: flights from Pago Pago, Samoa to Hawaii +2025-04-03 at 19:51:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 19:51:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:51:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 8 recovery crew and island or facility after USS Iwo Jima +2025-04-03 at 19:51:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 19:51:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ņ–ĐšĐŊĐžIwo Jima recovery crew destination after Pago Pago Samoa +2025-04-03 at 19:51:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 19:51:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: flights from Hickam Air Force Base, Hawaii to Pago Pago, Samoa +2025-04-03 at 19:51:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 19:51:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:51:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 8 recovery crew after USS Iwo Jima Samoa +2025-04-03 at 19:51:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 19:51:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:51:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "US mainland or island after USS Iwo Jima recovery for Apollo crew +2025-04-03 at 19:51:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +The Department of Defense provided recovery support commensurate with mission planning for Apollo l3. Because of the emergency which resulted in premature termination of the mission, additional support was provided by the Department of Defense and offers of assistance were made by many foreign nations, including England, France, Greece, Spain, Germany, Uruguay, Brazil, Kenya, the Netherlands, Nationalist China, and the Soviet Union. As a result of this voluntary support, a total of 2l ships and 17 aircraft were available for supporting an Indian Ocean landing, and 5l ships and 2l aircraft for an Atlantic Ocean landing. In the Pacific Ocean, there were l3 ships and 17 aircraft known to be available over and above the forces designated for primary recovery support. +------ + +2025-04-03 at 19:51:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:51:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "United States mainland after USS Iwo Jima recovery for Apollo 8 crew +2025-04-03 at 19:51:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 19:51:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:51:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 8 crew recovery from USS Iwo Jima flight not specified +2025-04-03 at 19:51:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 19:51:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:51:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "No answer +2025-04-03 at 19:51:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +structure during which the oxygen shelf was accidentally dropped with no apparent damage +------ + +2025-04-03 at 19:51:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:51:48 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:51:48 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:51:49 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_correctness:82 - Student lengths: [662, 9, 243, 143, 151, 370] +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [6, 6, 6, 6, 6, 6] +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_correctness:84 - Average student length: 263.00 +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 6.00 +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_correctness:86 - Length ratio: 43.83 +2025-04-03 at 19:51:49 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:51:49 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.492 Âą 0.301 +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.67 Âą 2.62 +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [1, 8, 0, 3, 3, 1] +2025-04-03 at 19:51:49 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations. +------ + +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +The Department of Defense provided recovery support commensurate with mission planning for Apollo l3. Because of the emergency which resulted in premature termination of the mission, additional support was provided by the Department of Defense and offers of assistance were made by many foreign nations, including England, France, Greece, Spain, Germany, Uruguay, Brazil, Kenya, the Netherlands, Nationalist China, and the Soviet Union. As a result of this voluntary support, a total of 2l ships and 17 aircraft were available for supporting an Indian Ocean landing, and 5l ships and 2l aircraft for an Atlantic Ocean landing. In the Pacific Ocean, there were l3 ships and 17 aircraft known to be available over and above the forces designated for primary recovery support. +------ + +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ + +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: ++21.0 +3.0 +------ +Result 2: +structure during which the oxygen shelf was accidentally dropped with no apparent damage +------ + +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 19:51:49 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The ship-based aircraft were deployed relative to the Iwo Jima and were on station 20 minutes prior to landing. They departed station to commence recovery activities upon receiving notice of visual contact with the descending command module. Figure l0.3-l depict an approximation of the recovery force positions just prior to the sighting of the command module. + + + +Figure l0.3-l.- Recovery support at earth landing. +------ +Result 2: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ + +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston. + +Upon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft. + +The following is a chronological listing of events during the recovery operations. +------ +Result 2: +Support for the primary recovery area consisted of the prime recovery ship, USS Iwo Jima, five helicopters from the Iwo Jima, and two HC-l30H rescue aircraft. Later, the experimental mine sweeper, USS Granville Hall, and two HC-l30H aircraft were added to the end-of-mission array. One of the helicopters, designated "Recovery," carried the flight surgeon, and was utilized for retrieval of the crew. Two of the helicopters, designated "Swim l" and "Swim 2," carried swimmers and the necessary recovery equipment. A fourth helicopter, designated "Photo" was used as a photographic platform, and the fifth helicopter, designated "Relay," served as a communications relay aircraft. The four aircraft, designated "Samoa Rescue l, 2, 3, and 4," were positioned to track the command module after exit from blackout, as well as to provide pararescue capability had the command module landed uprange or downrange of the target point. The USS Granville Hall was positioned to provide support in the event +------ + +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 5.0 +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.833 +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:51:49 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.133, Perfect scores: 0/6 +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.33 +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:51:49 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:51:49 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.318, Max reward: 0.631 +2025-04-03 at 19:51:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:51:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "breech plenum material for future spacecraft" +2025-04-03 at 19:51:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:51:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Material choices for low-out-gas toxicity and thermal management in spacecraft interior" +2025-04-03 at 19:51:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:51:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: non-toxic materials for spacecraft breech plenum liners +2025-04-03 at 19:51:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:51:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: radiation hardening coatings or materials for spacecraft breech plenum area +2025-04-03 at 19:51:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 19:51:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:51:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("breech plenum material for spacecraft" OR "breech plenum material on future spacecraft" OR "tunnel gusset protection on spacecraft") +2025-04-03 at 19:51:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:51:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Spacesuit cabin filter material" +2025-04-03 at 19:51:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:51:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: polyimide vs other materials for interior of breech plenum area in spacecraft +2025-04-03 at 19:51:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:51:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: polyimide material widespread adoption in spacecraft breech plenum area +2025-04-03 at 19:51:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:51:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:52:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("breech plenum material" OR "breech plenum seal material" OR "Polyimide seal material") +2025-04-03 at 19:52:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:52:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "spacesuit cabin dust filtration material" +2025-04-03 at 19:52:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ + +2025-04-03 at 19:52:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: polyimide material in spacecraft manufacturing industry +2025-04-03 at 19:52:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 19:52:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:52:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("O-ring vs Teflon ring" OR "NASA seal standards" OR "gusset seal requirements") +2025-04-03 at 19:52:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:52:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Nylon vacuum brushĪÎĩĪ‚ +2025-04-03 at 19:52:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Syneresis, or separation of a liquid from a solid, occurred in some of the canned sandwich spreads, particularly the ham salad. The free liquid escaped when the can was opened, and the salad was too dry to spread. The crew commented on the positive pressure in the bread packages, which was expected since there was only a slight vacuum on these packages. Any additional' vacuum would compress the bread.to an unacceptable state, and if the packages were punctured, the bread would become dry and hard. The crew recommended a.change which has been implemented wherein Velcro patches will be attached to the bread, mustard, and catsup packages. + +9.2.6 Radi ati on +------ + +2025-04-03 at 19:52:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: polyimide insulation in breech plenum area future spacecraft +2025-04-03 at 19:52:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:52:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:52:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("polyimide temperature range" OR "polyimide thermal stability" OR "polyimide vs extreme temperatures") +2025-04-03 at 19:52:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:52:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Nylon vacuum brush application in space dust filtration" +2025-04-03 at 19:52:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:52:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:52:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("breech plenum thermal conditions" OR "heat load on spacecraft components" OR "breech plenum material thermal performance") +2025-04-03 at 19:52:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:52:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Nylon brush application in spacesuit dust filtration" +2025-04-03 at 19:52:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 19:52:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:52:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("spacecraft material inspection report" OR " Apollo mission component material" OR "breech plenum material inspection") +2025-04-03 at 19:52:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 19:52:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "NASA spacesuit cabin dust filter material" +2025-04-03 at 19:52:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:52:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:52:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("polyimide for space applications" OR "polyimide in spacecraft engineering" OR "polyimide properties for extreme temperatures") +2025-04-03 at 19:52:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 19:52:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Battery material change from Teflon to Nylon" +2025-04-03 at 19:52:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +Evidence indicates that battery 2 may have experienced an electrical fault of some type. The most probable condition is electrolyte leaking from one or more cells and bridging the high-voltage or low-voltage terminal to the battery case (fig. 14-17). This bridging results in water electrolysis and subsequent ignition of the hydrogen and oxygen so generated. The accompanying "explosion" would then blow off or rupture the seal of the battery lid and cause both a thump and venting of the free liquids in the battery case, resulting in "snowflakes." + +Postflight tests have shown the following: +------ + +2025-04-03 at 19:52:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:52:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Electrolyte bridging in space battery potting material" +2025-04-03 at 19:52:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ + +2025-04-03 at 19:52:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:52:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "NASA spacesuit dust filter material" +2025-04-03 at 19:52:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:52:14 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:52:14 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:52:15 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, True, True] +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_correctness:82 - Student lengths: [158, 318, 214, 1927, 299, 564] +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [9, 9, 9, 9, 9, 9] +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_correctness:84 - Average student length: 580.00 +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 9.00 +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_correctness:86 - Length ratio: 64.44 +2025-04-03 at 19:52:15 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 19:52:15 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.454 Âą 0.361 +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 4.50 Âą 4.68 +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 8, 0, 13, 2, 4] +2025-04-03 at 19:52:15 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-03 at 19:52:15 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-03 at 19:52:15 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +Syneresis, or separation of a liquid from a solid, occurred in some of the canned sandwich spreads, particularly the ham salad. The free liquid escaped when the can was opened, and the salad was too dry to spread. The crew commented on the positive pressure in the bread packages, which was expected since there was only a slight vacuum on these packages. Any additional' vacuum would compress the bread.to an unacceptable state, and if the packages were punctured, the bread would become dry and hard. The crew recommended a.change which has been implemented wherein Velcro patches will be attached to the bread, mustard, and catsup packages. + +9.2.6 Radi ati on +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +Evidence indicates that battery 2 may have experienced an electrical fault of some type. The most probable condition is electrolyte leaking from one or more cells and bridging the high-voltage or low-voltage terminal to the battery case (fig. 14-17). This bridging results in water electrolysis and subsequent ignition of the hydrogen and oxygen so generated. The accompanying "explosion" would then blow off or rupture the seal of the battery lid and cause both a thump and venting of the free liquids in the battery case, resulting in "snowflakes." + +Postflight tests have shown the following: +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +a. Electrolyte can leak past the Teflon retention screens installe in each cell to prevent leakage. b.. The descent battery cells contain an excessive amount of free electrolyte. c. The potting does not adhere to the battery case, consequently, any free electrolyte can readily penetrate the interface between the potting and the case and bridge between the terminals and case. d. Once an electrolyte bridge is formed, electrolysis will produce hydrogen and oxygen gas. e. A bridge at the positive terminal can produce a current surge o: as much as l50 amperes. + +For Apollo l4 and subsequent missions, the descent batteries will be modified to minimize the hazards associated with electrolyte leakage. + + + +NASA-S-70-5859 + +Figure 14-l7.- Descent battery terminal configuration. +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:52:15 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\npropulsion system was changed to include a bypass line around the fuel/helium...', 'Result 1:\npropulsion system was changed to include a bypass line around the fuel/helium...', 'Result 1:\npropulsion system was changed to include a bypass line around the fuel/helium...', 'Result 1:\nwas added for cleaning the suits before ingress, the bristles on the vacuum b...', 'Result 1:\nwas added for cleaning the suits before ingress, the bristles on the vacuum b...', 'Result 1:\nwas added for cleaning the suits before ingress, the bristles on the vacuum b...', 'Result 1:\npropulsion system was changed to include a bypass line around the fuel/helium...', 'Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...', 'Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...', 'Result 1:\nThe extravehicular mobility unit underwent several modifications to improve l...'] +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ +Result 2: +During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units. +------ + +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 19:52:15 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.167, Perfect scores: 0/6 +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.33 +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:52:15 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:52:15 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.212, Max reward: 0.682 +2025-04-03 at 19:52:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:52:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what task was verified in a simulator before advising the crew to the Apollo astronauts +2025-04-03 at 19:52:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-03 at 19:52:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "task verification simulator before crew deployment " +2025-04-03 at 19:52:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ + +2025-04-03 at 19:52:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: task simulator mission verifier validator crew +2025-04-03 at 19:52:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 19:52:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space mission simulator task verification challenge or task validation +2025-04-03 at 19:52:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:52:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:52:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo mission simulator training procedures lunar module alignments and entry +2025-04-03 at 19:52:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:52:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apartment 11 mission simulator task verification +2025-04-03 at 19:52:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 19:52:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 1 command module post- flight testing +2025-04-03 at 19:52:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:52:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module undocking procedure verification simulator training Apollo 13 crew +2025-04-03 at 19:52:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:52:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:52:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: validate apollo 13 lunar landing and trans earth injection procedures +2025-04-03 at 19:52:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:52:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "simulator task verification Apollo mission crew prep routine +2025-04-03 at 19:52:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:52:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 1 task verified simulator training +2025-04-03 at 19:52:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:52:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 entry separation lunar module undocking training simulator +2025-04-03 at 19:52:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:52:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:52:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar landing and emergency plan +2025-04-03 at 19:52:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:52:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "abortion of apollo 11 mission procedure +2025-04-03 at 19:52:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +4.0 TRAJECTORY + +The planned trajectory profile was similar to that for Apollo 12 except for descent orbit insertion being performed with the service propulsion system and the targeting of the spent S--IvB stage for a lunar impact. The trajectory had been very close to the nominal flight plan up to the time of abort, which was the first in the Apollo program. Throughout the manned space program, techniques have been developed and tested for the real-tine determination of immediate abort requirements, but Apollo l3 presented the first situation in which their use was necesSary. Figure 3-l shows the mission profile, including the relative locations of all major maneuvers. +------ + +2025-04-03 at 19:52:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar landing trainer equipment +2025-04-03 at 19:52:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:52:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew pre-entry procedure verification launch mission +2025-04-03 at 19:52:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:52:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:52:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 mission simulator training +2025-04-03 at 19:52:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:52:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "abortion procedure lunar module simulation verification +2025-04-03 at 19:52:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:52:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 training for emergency situations +2025-04-03 at 19:52:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:52:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 MSC-02545 cryogenic oxygen tank anomaly report +2025-04-03 at 19:52:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:52:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:52:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar exploration crew preparedness for critical systems failures +2025-04-03 at 19:52:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 19:52:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 emergency procedures +2025-04-03 at 19:52:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:52:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:52:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module power-up and entry procedures +2025-04-03 at 19:52:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:52:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ØŦاŲ… mission problem solving +2025-04-03 at 19:52:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-03 at 19:52:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:52:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module systems power-up and assessment +2025-04-03 at 19:52:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:52:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 objectives +2025-04-03 at 19:52:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:52:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:52:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module systems re-entry preparation +2025-04-03 at 19:52:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:52:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:52:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: postanalysis of lunar module tunnel hatch +2025-04-03 at 19:52:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Hatch closeout in both spacecraft was normal, and a successful command module hatch integrity check was made, with a differential pressure of 3.4 psi. The command module environmental control and autopilot systems were activated, and the lunar module was undocked l hour before entry. Lunar module jettison was slightly louder than service module jettison and the lunar module was stable as it translated away using only tunnel pressure. While controllable by a single reaction control engine pulse, there was a. continuous pitch-up torque on the command module which persisted until entry. + +8.10 ENTRY AND LANDING +------ + +2025-04-03 at 19:52:40 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:52:40 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:52:40 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1531, 295, 687, 271, 230, 896] +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_correctness:84 - Average student length: 651.67 +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_correctness:86 - Length ratio: 65.17 +2025-04-03 at 19:52:40 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:52:40 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.554 Âą 0.426 +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 5.83 Âą 5.76 +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [17, 5, 8, 0, 0, 5] +2025-04-03 at 19:52:40 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +Hatch closeout in both spacecraft was normal, and a successful command module hatch integrity check was made, with a differential pressure of 3.4 psi. The command module environmental control and autopilot systems were activated, and the lunar module was undocked l hour before entry. Lunar module jettison was slightly louder than service module jettison and the lunar module was stable as it translated away using only tunnel pressure. While controllable by a single reaction control engine pulse, there was a. continuous pitch-up torque on the command module which persisted until entry. + +8.10 ENTRY AND LANDING +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +4.0 TRAJECTORY + +The planned trajectory profile was similar to that for Apollo 12 except for descent orbit insertion being performed with the service propulsion system and the targeting of the spent S--IvB stage for a lunar impact. The trajectory had been very close to the nominal flight plan up to the time of abort, which was the first in the Apollo program. Throughout the manned space program, techniques have been developed and tested for the real-tine determination of immediate abort requirements, but Apollo l3 presented the first situation in which their use was necesSary. Figure 3-l shows the mission profile, including the relative locations of all major maneuvers. +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:52:40 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ["Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nInitial outside observations through the lunar module windows indicated that ...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nThe Mission Control Center and the Manned Space Flight Network provided excel...', 'Result 1:\na. Perform selenological inspection, survey, and sampling of materials in a p...'] +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 19:52:40 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 19:52:40 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 19:52:40 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.133, Perfect scores: 0/6 +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.33 +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:52:40 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:52:40 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.496, Max reward: 0.789 +2025-04-03 at 19:52:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:52:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How long after pumpdown is pressure measured to verify vacuum integrity in jacketed vacuum vessels used in cryogenic applications? +2025-04-03 at 19:52:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 19:52:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how long after pumpdown is pressure in jacket measured to verify vacuum integrity +2025-04-03 at 19:52:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 19:52:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How long after pumpdown is the pressure in the jacket measured for vacuum integrity verification in cryogenic applications? +2025-04-03 at 19:52:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 19:52:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure in vacuum vessel gauge calibration +2025-04-03 at 19:52:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 19:52:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How long after pumpdown is pressure in the jacket measured to verify vacuum integrity in the context of vacuum insulation or cryogenic applications? +2025-04-03 at 19:52:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 19:52:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:52:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the standard time frame to measure vacuum integrity after catalyst pumpdown +2025-04-03 at 19:52:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 19:52:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA project Mercury or Apollo measurement intervals for vacuum integrity +2025-04-03 at 19:52:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 19:52:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Jacket pressure measurement timing post-pumpdown +2025-04-03 at 19:52:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 19:52:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Yale space program oxygen tank vacuum integrity measurement time +2025-04-03 at 19:52:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 19:52:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:52:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vacuum integrity measurement time after pumpdown +2025-04-03 at 19:52:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 19:52:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressuresensor measurement intervals for vacuum integrityResult 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 19:52:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 3-week vacuum jacket pressure measurement timing +2025-04-03 at 19:52:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 19:52:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Miniature cryogenic tank pressure measurement time post-pumpdown Airbus +2025-04-03 at 19:52:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 19:52:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:52:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vacuum integrity measurement post pumpdown human spaceflight +2025-04-03 at 19:52:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 19:52:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Suit jacket pressure verification timing range 2-3 weeks +2025-04-03 at 19:52:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 19:52:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Max duration for measuring jacket pressure after pumpdown for verifying vacuum integrity +2025-04-03 at 19:52:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 19:52:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:52:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the standard procedure for verifying vacuum integrity of pressurized systems on spacecraft after pumpdown +2025-04-03 at 19:52:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-03 at 19:52:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Typical minimum jacket pressure measurement timing after pumpdown +2025-04-03 at 19:52:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 19:52:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Effect of pumpdown duration on verifying vacuum integrity measurement +2025-04-03 at 19:52:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 19:52:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:52:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: is vacuum valve resistance part of verification of vacuum integrity +2025-04-03 at 19:52:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 19:52:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Typical minimum successful jacket pressure verification time post-pumpdown +2025-04-03 at 19:52:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 19:52:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Verification of vacuum integrity requirements for tank jacket pressure measurement +2025-04-03 at 19:52:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 19:52:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:53:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: standard test procedure for measuring vacuum system integrity +2025-04-03 at 19:53:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 19:53:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Typical minimum successful jacket pressure measurement time after pumpdown +2025-04-03 at 19:53:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 19:53:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ASCO or NASA document on vacuum integrity measurement after pumpdown +2025-04-03 at 19:53:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 19:53:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:53:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: measurement of vacuum system pressure changes after vacuum pumpdown +2025-04-03 at 19:53:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-03 at 19:53:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Jacket pressure measurement post-pumpdown ground verification +2025-04-03 at 19:53:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +according to ground instructions. +------ + +2025-04-03 at 19:53:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA technical standard for measuring vacuum integrity in spacecraft +2025-04-03 at 19:53:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 19:53:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:53:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ground confirmed jacket pressure measurement timing post-pumpdown +2025-04-03 at 19:53:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 19:53:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA requirements for verifying vacuum integrity of spacecraft jacket pressure after pumpdown +2025-04-03 at 19:53:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 19:53:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:53:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: jackets pressure measurement timing after pumping 59 hours +2025-04-03 at 19:53:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 19:53:07 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:53:07 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:53:07 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, False, False] +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_correctness:82 - Student lengths: [599, 357, 282, 945, 1690, 1854] +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [12, 12, 12, 12, 12, 12] +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_correctness:84 - Average student length: 954.50 +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 12.00 +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_correctness:86 - Length ratio: 79.54 +2025-04-03 at 19:53:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-03 at 19:53:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.642 Âą 0.361 +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 5.17 Âą 3.98 +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [8, 3, 1, 0, 10, 9] +2025-04-03 at 19:53:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle. + +The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-03 at 19:53:07 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +according to ground instructions. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +Cryogenic storage system operation was satisfactory until 46:40:09, when the quantity indication was lost for oxygen tank 2 (section 14.l.l). At about 56 hours , the pressure in oxygen tank 2 suddenly dropped to zero and the pressure in oxygen tank l began to decay until all primary oxygen was lost. As a result, power was lost from fuel cells l and 3, and after Oxygen was essentially depleted from tank l, fuel cell 2 was taken offline. After the flight, a comprehensive review of the history of cryogenic oxygen tank 2 was made to determine wheth.er an unfavorable condition could have existed prior to laumch. This review included test records, materials review dispositions, and failure reports. No positive indication of any unfavorable conditions prior to shipment to the launch site could be found in the testing or inspections conducted. However, to accomplish a modification on the vac-ion pumps, the complete oxygen shelf, including the oxygen tanks, was removed from the service module +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2). + + + +(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement ¡ operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 5.0 +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.833 +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:53:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.133, Perfect scores: 0/6 +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.33 +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:53:07 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:53:07 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.458, Max reward: 0.688 +2025-04-03 at 19:53:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:53:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the temperature range of an antenna in passive thermal control mode? +2025-04-03 at 19:53:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-03 at 19:53:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "NASA Orion spacecraft antenna temperature range passive thermal control mode +2025-04-03 at 19:53:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-03 at 19:53:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: thermal control mode antenna temperature range +2025-04-03 at 19:53:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 19:53:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "spacecraft passive thermal control system antenna operational temperature range" +2025-04-03 at 19:53:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-03 at 19:53:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:53:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the temperature range of the antenna in space during passive thermal control mode? +2025-04-03 at 19:53:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-03 at 19:53:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature range windows of Orion spacecraft antenna passive thermal control mode +2025-04-03 at 19:53:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-03 at 19:53:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo spacecraft antennae temperature control MOS temperate range +2025-04-03 at 19:53:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:53:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:53:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Antenna temperature range in space S-band steerable antenna passive thermal control mode +2025-04-03 at 19:53:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-03 at 19:53:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Orion spacecraft antenna thermal characteristics +2025-04-03 at 19:53:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 19:53:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: sandwich antenna temperature control +2025-04-03 at 19:53:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +------ + +2025-04-03 at 19:53:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:53:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: specific action taken during passive thermal control mode for antenna temperature stabilization +2025-04-03 at 19:53:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-03 at 19:53:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-band antenna temperature range Orion passive thermal control mode +2025-04-03 at 19:53:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 19:53:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo s-band antenna passive thermal control temperature range scanning mode +2025-04-03 at 19:53:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 19:53:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:53:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-band antenna temperature control system Orion +2025-04-03 at 19:53:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ + +2025-04-03 at 19:53:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo passive thermal control system attitude maneuvers +2025-04-03 at 19:53:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-03 at 19:53:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:53:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-band antenna temperature range Orion passive mode +2025-04-03 at 19:53:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ + +2025-04-03 at 19:53:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo passive thermal control mode yaw maneuver schedule +2025-04-03 at 19:53:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 19:53:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:53:25 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:53:25 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:53:25 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, True, False, False, False, False] +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_correctness:82 - Student lengths: [129, 300, 603, 577, 808, 111] +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [29, 29, 29, 29, 29, 29] +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_correctness:84 - Average student length: 421.33 +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 29.00 +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_correctness:86 - Length ratio: 14.53 +2025-04-03 at 19:53:25 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:53:25 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.529 Âą 0.406 +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 3.83 Âą 2.85 +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [4, 6, 6, 0, 0, 7] +2025-04-03 at 19:53:25 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ + +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ + +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual +------ + +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ +Result 2: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ + +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +After the free-returm midcourse correction, the spacecraft was maneuvered manually to the passive thermal control mode attitudes. The passive thermal control mode techniques consisted of maneuvering in the pulse mode 9o degrees in yaw once each hour using the pulse mode. To conserve power, the attitude indicators were turned off after the initial passive thermal control mode was started, and attitude monitoring was accomplished by observing gimbal angle readouts from the display keyboard. +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 19:53:25 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 19:53:25 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:53:25 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.167, Perfect scores: 0/6 +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.00 +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:53:25 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:53:25 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.277, Max reward: 0.663 +2025-04-03 at 19:53:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:53:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the meaning of "first number in the list" in different contexts, such as ordering, sequences, or other mathematical operations? +2025-04-03 at 19:53:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +4.0 TRAJECTORY + +The planned trajectory profile was similar to that for Apollo 12 except for descent orbit insertion being performed with the service propulsion system and the targeting of the spent S--IvB stage for a lunar impact. The trajectory had been very close to the nominal flight plan up to the time of abort, which was the first in the Apollo program. Throughout the manned space program, techniques have been developed and tested for the real-tine determination of immediate abort requirements, but Apollo l3 presented the first situation in which their use was necesSary. Figure 3-l shows the mission profile, including the relative locations of all major maneuvers. +------ + +2025-04-03 at 19:53:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"What is the first number in a list of natural numbers?" +"What is the first item in an ordered list?" +"Pattern number one in a list sequence?" +2025-04-03 at 19:53:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:53:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the first number in a standard numerical list? +2025-04-03 at 19:53:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:53:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first ranking in top teams in the NFL +2025-04-03 at 19:53:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +boost phases were within acceptable spacecraft structural. design limits. +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 19:53:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first number in list +2025-04-03 at 19:53:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:53:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is a numerical list of natural, fundamental quantities, fundamental sequence? +2025-04-03 at 19:53:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:53:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:53:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the numbers associated with astronauts at the beginning of the provided text, specifically the "1038.6", "0.5", or the "4.0"? +2025-04-03 at 19:53:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:53:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: list of first numbers +2025-04-03 at 19:53:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:53:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first value of "X Z" in data +2025-04-03 at 19:53:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +-1.2 40.4 +0.4 +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 19:53:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the first number in the list +2025-04-03 at 19:53:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:53:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Is the first number in Table D-I the numerical value in a sequence or pattern, such as a list of data from a system performance analysis? +2025-04-03 at 19:53:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 19:53:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:53:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: The numbers 1038.6 and 0.5 and other numbers mentioned in NASA documents from 1970, specifically those related to manned spacecraft and trajectory execution. +2025-04-03 at 19:53:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:53:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "First" + " initial" +2025-04-03 at 19:53:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: ++21.0 +3.0 +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 19:53:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the numerical value of the first data point in Table D-I, specifically the coefficient in the X-axis data? +2025-04-03 at 19:53:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 19:53:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:53:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first event in translunar phase +2025-04-03 at 19:53:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 19:53:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:53:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first time > translunar injection +2025-04-03 at 19:53:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 19:53:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:53:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first translunar injection time +2025-04-03 at 19:53:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 19:53:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:53:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first translunar injection + 6 hours +2025-04-03 at 19:53:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 19:53:44 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:53:44 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:53:44 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, False, True, True] +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_correctness:82 - Student lengths: [486, 256, 232, 1586, 20, 367] +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [6, 6, 6, 6, 6, 6] +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_correctness:84 - Average student length: 491.17 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 6.00 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_correctness:86 - Length ratio: 81.86 +2025-04-03 at 19:53:44 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 19:53:44 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.571 Âą 0.191 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 5.50 Âą 5.91 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [3, 1, 2, 7, 2, 18] +2025-04-03 at 19:53:44 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +4.0 TRAJECTORY + +The planned trajectory profile was similar to that for Apollo 12 except for descent orbit insertion being performed with the service propulsion system and the targeting of the spent S--IvB stage for a lunar impact. The trajectory had been very close to the nominal flight plan up to the time of abort, which was the first in the Apollo program. Throughout the manned space program, techniques have been developed and tested for the real-tine determination of immediate abort requirements, but Apollo l3 presented the first situation in which their use was necesSary. Figure 3-l shows the mission profile, including the relative locations of all major maneuvers. +------ + +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +boost phases were within acceptable spacecraft structural. design limits. +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +-1.2 40.4 +0.4 +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: ++21.0 +3.0 +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 19:53:44 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nboost phases were within acceptable spacecraft structural. design limits.\n---...', 'Result 1:\n-1.2 40.4 +0.4\n------\nResult 2:\nEvent Weight, 1b Center of gravity, in. Momen...', 'Result 1:\n+21.0 +3.0\n------\nResult 2:\nTranslunar phase Event Reference body Time, hr:mi...', 'Result 1:\nTranslunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitud...', 'Result 1:\nTranslunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitud...', 'Result 1:\nFollowing translunar injection, earth weather photography was conducted for a...', 'Result 1:\nFollowing translunar injection, earth weather photography was conducted for a...'] +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ + +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 19:53:44 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', 'Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', 'Result 1:\nupdated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 ...'] +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:53:44 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.133, Perfect scores: 0/6 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.17 +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 19:53:44 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:53:44 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.634, Max reward: 0.916 +2025-04-03 at 19:53:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:53:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +percentage of damaged stitches in roller blinds due to manufacturing defects or wear and tear +2025-04-03 at 19:53:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:53:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What percentage of stitch holes in a typical window shade have cracks extending from them? +2025-04-03 at 19:53:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:53:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: statistics on stitch hole defects in window shades +2025-04-03 at 19:53:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:53:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:53:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +percentage of stitch holes cracked in window shades due to manufacturing defects +2025-04-03 at 19:53:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:53:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: percentage of stitch holes with cracks in Apollo 13 left-hand window shade +2025-04-03 at 19:53:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:53:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:53:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +percentage of stitches with cracks in window shades due to manufacturing defects +2025-04-03 at 19:53:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:53:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Aporia 13 window shade stitching failure percentage +2025-04-03 at 19:53:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:53:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:53:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +percentage of stitches with cracks in lunar module window shades +2025-04-03 at 19:53:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:53:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 window shade stitching failure < 80% +2025-04-03 at 19:53:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:53:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:53:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +cracked stitches in lunar module window shades percentage of total stitches +2025-04-03 at 19:53:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:53:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Aporia 13 80% stitching failure window shade cracks +2025-04-03 at 19:53:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:53:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:53:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Aporia 13 window shade < 80% stitching failure percentage +2025-04-03 at 19:53:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:53:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:54:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 window shade 80% stitching failure <1/8 inch cracks +2025-04-03 at 19:54:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:54:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:54:01 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:54:01 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:54:01 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, True, False] +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_correctness:82 - Student lengths: [373, 275, 304, 509, 166, 483] +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_correctness:84 - Average student length: 351.67 +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_correctness:86 - Length ratio: 35.17 +2025-04-03 at 19:54:01 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:54:01 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.375 Âą 0.425 +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.17 Âą 2.79 +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 0, 5, 1, 7, 0] +2025-04-03 at 19:54:01 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 19:54:01 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 19:54:01 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ +Result 2: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ + +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +------ +Result 2: +The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight. + +This anomaly is closed. + +14.3 GOVERNMENT FURNISHED EQUIPMENT + +14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera + +For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies. +------ + +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves. + +This anomaly is closed. + +14.2.5 Cracked Window Shade + +The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade. + + + +Figure 14-21.- Cracked left-hand window shade. + +Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight. +2025-04-03 at 19:54:01 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 19:54:01 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.067, Perfect scores: 0/6 +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.17 +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:54:01 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:54:01 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.225, Max reward: 0.688 +2025-04-03 at 19:54:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:54:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: "service propulsion system fuel consumption SpaceX" +2025-04-03 at 19:54:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-03 at 19:54:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Saturn V rocket service propulsion system fuel consumption" +2025-04-03 at 19:54:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-03 at 19:54:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 SPS attitude control delta-v" +2025-04-03 at 19:54:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:54:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Service propulsion system fuel consumption" +2025-04-03 at 19:54:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-03 at 19:54:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:54:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 Saturn V rocket fuel consumption" +2025-04-03 at 19:54:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:54:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 SPS consumption fuel" +2025-04-03 at 19:54:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 19:54:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:54:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Saturn V total fuel consumption" +2025-04-03 at 19:54:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 19:54:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo SPS propellant consumption mass" +2025-04-03 at 19:54:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:54:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:54:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 Saturn V mission payload status" +2025-04-03 at 19:54:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:54:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 Service Propulsion System hydrogen consumption" +2025-04-03 at 19:54:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 19:54:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:54:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 mission fuel consumption data" +2025-04-03 at 19:54:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:54:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "hydrogen fuel load increase Apollo 11" +2025-04-03 at 19:54:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:54:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:54:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Total fuel consumption Apollo 11 Saturn V mission" +2025-04-03 at 19:54:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:54:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 excursion logs Service Propulsion System" +2025-04-03 at 19:54:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:54:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:54:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Saturn V rocket fuel consumption" +2025-04-03 at 19:54:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 19:54:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 12 Service Propulsion System fuel consumption" +2025-04-03 at 19:54:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:54:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:54:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Saturn V main stage fuel consumption" +2025-04-03 at 19:54:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ + +2025-04-03 at 19:54:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:54:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Saturn V main stage fuel consumption" +2025-04-03 at 19:54:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ + +2025-04-03 at 19:54:26 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:54:26 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:54:26 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_correctness:82 - Student lengths: [74, 485, 1696, 1886, 67, 449] +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_correctness:84 - Average student length: 776.17 +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 4.00 +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_correctness:86 - Length ratio: 194.04 +2025-04-03 at 19:54:26 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:54:26 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.354 Âą 0.322 +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 3.67 Âą 4.53 +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [1, 0, 9, 11, 0, 1] +2025-04-03 at 19:54:26 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 19:54:26 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ + +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +If the Saturn V vehicle is charged to a potential of several million volts, corona discharges will be produced which, in turn, generate radio noise. The onset of these discharges should occur very soon after liftoff and reach a maximum when the launch vehicle is still close to the ground. Radio noise records strongly support this conclusion. The sudden onset of the noise probably corresponds closely to lift-off. It is interesting that, at about 15 seconds after lift-off, the noise became enhanced at the lower rather than the higher frequencies. This phenomenon implies that larger discharges occur at these times. The most intense discharges would be expected to occur soon after the launch vehicle and its exhaust plume clear the launch tower. + +1l.l.3 Measurement of Pelluric Current +------ + +2025-04-03 at 19:54:26 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nService module.- At the time the system was powered down, reaction control sy...', 'Result 1:\nFuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nTotal oxygen usage from the three lunar module oxygen tanks was 20.3 pounds o...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...'] +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +Fuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6 + +7.2.2 Reaction Control Propellants + +The reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature. + +Fuel, lb Oxidi zerīŧŒ lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5 Total 166 + +7.2.3 0xygen + +Actual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data. + +Loaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4 +------ + +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:54:26 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nFuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5...', 'Result 1:\nFuel, 1b Oxi di zerīŧŒ lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5...', 'Result 1:\nThe command module was completely powered down at 58 hours 40 minutes, at whi...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nSupplement number Title Publication date/status Apollo 12 Trajectory Reconstr...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...'] +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 19:54:26 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:54:26 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:54:26 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:54:26 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.067, Max reward: 0.200 +2025-04-03 at 19:54:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:54:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happened at 32:21:49? minimum impulse engine firing - all systems +2025-04-03 at 19:54:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 19:54:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "What triggered unplanned minimum impulse engine firing at 32:21:49" +2025-04-03 at 19:54:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 19:54:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: minimum impulse engine firing 32:21:49 +2025-04-03 at 19:54:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 19:54:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "unplanned minimum impulse engine firing timestamp 32:21:49" +2025-04-03 at 19:54:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ + +2025-04-03 at 19:54:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:54:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft unplanned impulse engine firing Transearth injection +2025-04-03 at 19:54:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 19:54:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("unplanned minimum impulse engine firing" AND "32:21:49" AND " Space Shuttle" +2025-04-03 at 19:54:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-03 at 19:54:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "NASA Apollo 9 unplanned minimum impulse engine firing" +2025-04-03 at 19:54:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 19:54:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:54:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: computer-controlled spacecraft transearth injection details +2025-04-03 at 19:54:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-03 at 19:54:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("Apollo 13 mission aborted due to tank failure" AND "32:21:49" AND "cryogenic oxygen tank" +2025-04-03 at 19:54:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:54:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 9 engine malfunction 32:21:49" +2025-04-03 at 19:54:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 19:54:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:54:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system uses +2025-04-03 at 19:54:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:54:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("Apollo 13 mission" AND "hydrogen pressure" AND "low pressure" AND "56 hours") +2025-04-03 at 19:54:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:54:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:54:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo transearth injection unplanned minimum impulse firing analysis +2025-04-03 at 19:54:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-03 at 19:54:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("Apollo 13 training" AND "launch preparedness" AND "hydrogen pressure concerns" +2025-04-03 at 19:54:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:54:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:54:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection with limited range data power amplifier off +2025-04-03 at 19:54:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 19:54:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("Houston training site" AND "Apollo 13" AND "Commander lunar landing training vehicle") +2025-04-03 at 19:54:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:54:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:54:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 transearth injection communication power management +2025-04-03 at 19:54:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:54:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("Apollo 12" AND "Apollo 13" AND "same mission" +2025-04-03 at 19:54:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:54:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:54:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("Apollo 12 mission crew" AND "Apollo 13 mission" +2025-04-03 at 19:54:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:54:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:54:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("Charles C. Fuller" AND "Apollo 12" +2025-04-03 at 19:54:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:54:47 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:54:47 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:54:47 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, False, False] +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_correctness:82 - Student lengths: [754, 1957, 227, 220, 522, 58] +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [112, 112, 112, 112, 112, 112] +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_correctness:84 - Average student length: 623.00 +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 112.00 +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_correctness:86 - Length ratio: 5.56 +2025-04-03 at 19:54:47 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:54:47 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.483 Âą 0.398 +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 3.33 Âą 3.50 +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [7, 9, 1, 0, 3, 0] +2025-04-03 at 19:54:47 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control. +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ +Result 2: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Because of a sudden loss of pressure at approximately 56 hours from one of the two service module cryogenic oxygen tanks in bay 4, primary electrical power was lost and the mission was aborted. Therefore, the remainder of this section will consider only the abort profile, since the trajectory prior to the tank incident was nearly identical to that of Apollo l2, including the first midcourse maneuver to a non-free-return profile, as shown in figure 3-l. The major trajectory difference from Apollo l2 resulted from an early shutdown of the center engine in the S-II stage of the Saturn V, the subsequent staging and insertion times were somewhat later than planned. A listing of significant mission events is contained in table 3-I. + + + +Figure 3-l.- Apollo l3 mission profile. + +TABLE 3-I.- SEQUENCE OF EVENTS +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-03 at 19:54:47 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly. + +This anomaly is closed. + +14.3.2 Failure of the Interval Timer Set Knob + +The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin. +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ + +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +2025-04-03 at 19:54:47 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:54:47 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.83 +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:54:47 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:54:47 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.179, Max reward: 0.711 +2025-04-03 at 19:54:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:54:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: velocity signal S-IVB seismic station Impact +2025-04-03 at 19:54:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 19:54:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 S-IVB post impact seismic station velocity +2025-04-03 at 19:54:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:54:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Skylab S-IVB impact location and distance to seismic station" +2025-04-03 at 19:54:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 19:54:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:54:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: velocity at which the initial signal from the S-IVB impact travelled to the seismic station +2025-04-03 at 19:54:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:54:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 S-IVB impact seismic velocity +2025-04-03 at 19:54:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:54:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"What is the velocity of an object 125 miles from a seismic station after 17 minutes?" +2025-04-03 at 19:54:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 19:54:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:54:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission S-IVB impact initial signal velocity +2025-04-03 at 19:54:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 19:54:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"velocity of seismic signal to earth from moon's surface at 125 miles" +2025-04-03 at 19:54:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:54:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:54:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 S-IVB impact velocity +2025-04-03 at 19:54:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:54:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"4.8 km/sec distance in miles" +2025-04-03 at 19:54:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +TABLE 4-II.- TRAJECTORY PARAMETERS +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 19:54:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:55:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 model S-IVB impact velocity calculation +2025-04-03 at 19:55:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:55:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:55:02 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:55:02 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:55:02 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, False, False, False] +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_correctness:82 - Student lengths: [68, 344, 121, 819, 242, 953] +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_correctness:84 - Average student length: 424.50 +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_correctness:86 - Length ratio: 42.45 +2025-04-03 at 19:55:02 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:55:02 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.375 Âą 0.398 +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.83 Âą 2.03 +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [5, 0, 2, 4, 0, 0] +2025-04-03 at 19:55:02 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-03 at 19:55:02 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ + +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +TABLE 4-II.- TRAJECTORY PARAMETERS +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-03 at 19:55:02 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon. +2025-04-03 at 19:55:02 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 19:55:02 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.50 +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:55:02 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:55:02 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.198, Max reward: 0.724 +2025-04-03 at 19:55:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:55:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How many seconds late orbital insertion occurred in space +2025-04-03 at 19:55:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 19:55:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: in ako 2018 insoy spectrography orbited and inserted planet mars +2025-04-03 at 19:55:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 19:55:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:55:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: entry and landing mars ranger 7 +2025-04-03 at 19:55:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 19:55:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:55:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: power-up of ranger 7 before entry and landing +2025-04-03 at 19:55:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 19:55:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:55:11 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:55:11 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:55:11 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_correctness:82 - Student lengths: [766, 52, 458, 98, 1102, 370] +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_correctness:84 - Average student length: 474.33 +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 5.00 +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_correctness:86 - Length ratio: 94.87 +2025-04-03 at 19:55:11 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:55:11 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.167 Âą 0.251 +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 0.67 Âą 1.11 +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 1, 0, 0, 0, 3] +2025-04-03 at 19:55:11 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 19:55:11 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 19:55:11 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe planned launch and earth parking orbit phases for this mission were very ...'] +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 19:55:11 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 19:55:11 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 19:55:11 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 19:55:11 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...', 'Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...', 'Result 1:\nThe unprecedented powered-down state of the command module required generatio...'] +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:55:11 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.067, Perfect scores: 0/6 +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.17 +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:55:11 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:55:11 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.156, Max reward: 0.737 +2025-04-03 at 19:55:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:55:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what unit is used to measure altitude above lunar surface +2025-04-03 at 19:55:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 19:55:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar altitude measurement" +2025-04-03 at 19:55:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ + +2025-04-03 at 19:55:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"lunar altitude measurement unit +2025-04-03 at 19:55:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 19:55:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar context altitude unit +2025-04-03 at 19:55:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:55:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar altitude unit +2025-04-03 at 19:55:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 19:55:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar altitude unit" +2025-04-03 at 19:55:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 19:55:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:55:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: altitude unit lunar surface +2025-04-03 at 19:55:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-03 at 19:55:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar mean radius" +2025-04-03 at 19:55:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 19:55:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar altitude unit definition +2025-04-03 at 19:55:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 19:55:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar altitude unit lmpdriver Apollo" +2025-04-03 at 19:55:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:55:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:55:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar radius formula and value" +2025-04-03 at 19:55:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 19:55:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar Module Flight-Path Angle unit +2025-04-03 at 19:55:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 19:55:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 10 spacecraft configuration" +2025-04-03 at 19:55:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:55:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:55:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar mean radius formula" +2025-04-03 at 19:55:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 19:55:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar altitude measurement Apollo +2025-04-03 at 19:55:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 19:55:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo l0 mission analysis navigation communication systems" +2025-04-03 at 19:55:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:55:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:55:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "mean lunar radius formula physics" +2025-04-03 at 19:55:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ + +2025-04-03 at 19:55:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar surface altitude measurement Apollo +2025-04-03 at 19:55:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 19:55:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 9 navigation system files" +2025-04-03 at 19:55:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:55:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:55:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar mean radius math formula" +2025-04-03 at 19:55:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 19:55:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar surface altitude measurement Apollo 13 +2025-04-03 at 19:55:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:55:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 9 navigation algorithms" +2025-04-03 at 19:55:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:55:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:55:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar mean radius units of measurement" +2025-04-03 at 19:55:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 19:55:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar altitude measurement altitude above lunar surface Apollo 13 experiment +2025-04-03 at 19:55:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:55:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar altitude above earth" +2025-04-03 at 19:55:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 19:55:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:55:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no relevant information +2025-04-03 at 19:55:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 19:55:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar surface thermal conductivity experiment unit +2025-04-03 at 19:55:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 19:55:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar altitude reference unit" +2025-04-03 at 19:55:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 19:55:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:55:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no relevant information +2025-04-03 at 19:55:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 19:55:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: heat flow experiment altimetry unit +2025-04-03 at 19:55:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-03 at 19:55:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "accelerometer bias compensation" +2025-04-03 at 19:55:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:55:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:55:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module altitude measurement" +2025-04-03 at 19:55:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 19:55:40 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:55:40 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:55:40 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, True, False, False] +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_correctness:82 - Student lengths: [221, 1032, 320, 724, 450, 1806] +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [13, 13, 13, 13, 13, 13] +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_correctness:84 - Average student length: 758.83 +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 13.00 +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_correctness:86 - Length ratio: 58.37 +2025-04-03 at 19:55:40 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 19:55:40 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.596 Âą 0.257 +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 14.33 Âą 22.48 +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [2, 64, 1, 9, 1, 9] +2025-04-03 at 19:55:40 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +The charged particle lunar environment experiment was designed to measure the energy of protons and electrons in the energy range of 4o to 70 electron volts. The experiment consisted of two detector/analyzer packages, each oriented for minimum exposure to the eclystic path of the sun, one for the east-west plane and one for the north-south plane. Each of the detector packages had six particle energy detectors. A complete measurement of all energy ranges would be made every 19.4 seconds. + +A.3.3 Cold Cathode Gage Experiment + +The cold cathode gage experiment was designed to measure the density of the lunar atmosphere by sensing the particle density immediately around its deployed position. An electrical current would be produced in the gage proportional to particle density. Pressure of the ambient atmosphere could be calculated, based on the measurements of the density of the neutralatoms. +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The lunar module platform was coarse aligned to the command module platform a few hours after the oxygen tank incident in preparation for the midcourse correction to enter a free-return trajectory. In preparing for the transearth injection maneuver, a check of the platform alignment accuracy was completed by letting the computer point the alignment optical telescope at the sun as though marks were to be taken. Results of the sun check angles indicated a platform misalignments about any axis of approximately half the allowable l-degree limit; therefore, a platform realignment was not required before the maneuver. +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ +Result 2: +The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13. + +The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay. + +NASA-S-70-5864 + + + +Figure A-l.- Experiment subpackage number l. + +A.3.1 Heat Flow Experiment + +The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements. +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 19:55:40 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nA.2 LUNAR MODULE\n------\nResult 2:\nAfter the oxygen tank incident, the platfor...'] +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersectīŧŒfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 5.0 +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.833 +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:55:40 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.233, Perfect scores: 0/6 +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.67 +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 19:55:40 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:55:40 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.223, Max reward: 0.581 +2025-04-03 at 19:55:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:55:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Antenna heater temperature Hyundai Tucson 2020 Corning Gorilla Glass" +2025-04-03 at 19:55:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:55:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: antenna temperature after heaters off +2025-04-03 at 19:55:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:55:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature of antenna heaters turned off temperature cell phone signal +2025-04-03 at 19:55:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 19:55:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:55:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Antenna temperature temperature control temperature -67°C" +2025-04-03 at 19:55:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 19:55:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature of antenna actors in space after heaters off +2025-04-03 at 19:55:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:55:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: antenna temperature during power off conditions S-band communication +2025-04-03 at 19:55:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-03 at 19:55:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:55:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Antenna heater temperature range for communication satellites" +2025-04-03 at 19:55:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 19:55:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: antenna temperature range after heaters off +2025-04-03 at 19:55:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 19:55:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:55:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"S-band antenna operating temperature" +2025-04-03 at 19:55:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ + +2025-04-03 at 19:55:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fans or heaters on high-gain antenna +2025-04-03 at 19:55:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 19:55:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:55:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: relevance of antenna heaters in high-gain antenna radio astronomy +2025-04-03 at 19:55:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 19:55:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:55:54 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:55:54 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:55:54 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_correctness:82 - Student lengths: [306, 134, 350, 1263, 357, 348] +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [11, 11, 11, 11, 11, 11] +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_correctness:84 - Average student length: 459.67 +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 11.00 +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_correctness:86 - Length ratio: 41.79 +2025-04-03 at 19:55:54 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_format:228 - Responses ending properly: 0/6 +2025-04-03 at 19:55:54 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.375 Âą 0.398 +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.83 Âą 2.03 +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [4, 5, 0, 0, 2, 0] +2025-04-03 at 19:55:54 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ + +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ + +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +14.l.4 High-Gain Antenna Acquisition Problem + +Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station. +------ +Result 2: +When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10. +------ + +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 19:55:54 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 19:55:54 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +------ +Result 2: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ + +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +6.3 COMMUNICATIONS EQUIPMENT + +S-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality. +------ +Result 2: +At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data. + +From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery. + +5.5 INS TRUMENTATION +------ + +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66° F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\circ}$ F。 + +6.4 GUIDANCE, NAVIGATION AND CONTROL + +System performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit. + +6.4.1 Attitude Control +2025-04-03 at 19:55:54 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 19:55:54 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.033, Perfect scores: 0/6 +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:55:54 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:55:54 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.221, Max reward: 0.741 +2025-04-03 at 19:55:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:56:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: rise rate from liftoff to first descent maneuver in orbital vehicles +2025-04-03 at 19:56:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 19:56:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: arthimis mission ascent propulsion average rise rate +2025-04-03 at 19:56:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 19:56:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: the average rise rate from lift-off to first descent propulsion maneuver for a specific space mission +2025-04-03 at 19:56:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 19:56:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Space Shuttle ascent and descent rate" +2025-04-03 at 19:56:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 19:56:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:56:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: descent propulsion maneuver duration from table 4-Il +2025-04-03 at 19:56:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:56:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: artemis mission ascent to descent propulsion rates +2025-04-03 at 19:56:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:56:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 transearth injection +2025-04-03 at 19:56:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:56:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Space Shuttle ascent to orbit velocity change" +2025-04-03 at 19:56:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 19:56:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:56:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: descent propulsion system instart time to engine shutdown for transearth injection +2025-04-03 at 19:56:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 19:56:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: artemis mission ascent to free-return trajectory +2025-04-03 at 19:56:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ + +2025-04-03 at 19:56:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 transearth descent propulsion average rise rate +2025-04-03 at 19:56:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:56:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Space Shuttle descent and ascent rate of change" +2025-04-03 at 19:56:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 19:56:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:56:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first transearth injection descent propulsion system duration +2025-04-03 at 19:56:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 19:56:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: artemis mission midcourse correction rise rate +2025-04-03 at 19:56:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:56:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 transearth injection rise rate +2025-04-03 at 19:56:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:56:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Space Shuttle optimal ascent time to first engine burn" +2025-04-03 at 19:56:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 19:56:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:56:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 spacecraft lift-off to midcourse correction rise rate +2025-04-03 at 19:56:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:56:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 first transearth rise rate +2025-04-03 at 19:56:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:56:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Space Shuttle ascent and descent propulsion system performance" +2025-04-03 at 19:56:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:56:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:56:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 ascent phase rise rate +2025-04-03 at 19:56:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 19:56:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Space Shuttle ascent and descent propulsion system capabilities" +2025-04-03 at 19:56:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:56:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:56:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 ascent midcourse rise rate +2025-04-03 at 19:56:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:56:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Space Shuttle ascent velocity profile" +2025-04-03 at 19:56:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 19:56:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:56:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo command and lunar module ascent rise rate +2025-04-03 at 19:56:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:56:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Space Shuttle lift-off to first main engine cutoff rate" +2025-04-03 at 19:56:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 19:56:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:56:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module ascent propulsion rise rate +2025-04-03 at 19:56:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 19:56:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Space Shuttle ascent velocity profile (numerical data)" +2025-04-03 at 19:56:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 19:56:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:56:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module ascent stage rise rate +2025-04-03 at 19:56:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:56:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Space Shuttle ascent and separation velocities" +2025-04-03 at 19:56:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 19:56:20 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:56:20 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:56:20 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 2/6 answers correct +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_correctness:82 - Student lengths: [231, 1838, 371, 1103, 187, 1414] +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_correctness:84 - Average student length: 857.33 +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_correctness:86 - Length ratio: 85.73 +2025-04-03 at 19:56:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:56:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.608 Âą 0.433 +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 4.83 Âą 4.10 +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [4, 0, 0, 10, 5, 10] +2025-04-03 at 19:56:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46 +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 19:56:20 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nThe vehicle was launched on an azimuth 90 degrees east of north, and a roll m...', 'Result 1:\nManeuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...'] +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 19:56:20 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 19:56:20 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:56:20 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...', 'Result 1:\nThe accuracy of the first midcourse correction (table 4-Il), which placed the...', 'Result 1:\nThe accuracy of the first midcourse correction (table 4-Il), which placed the...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...', 'Result 1:\nSeismic signals were first recorded 28.4 seconds after impact and continued f...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...'] +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:56:20 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...'] +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The accuracy of the first midcourse correction (table 4-Il), which placed the spacecraft on the non-free-return trajectory, was such that a maneuver was not required at the third planned option point. However, because of the oxygen tank incident, a 38-ft/sec midcourse maneuver was performed at 6l:29:44 using the descent engine to return the spacecraft to a free-return trajectory. This maneuver alone would have caused the command module to nominally land in the Indian Ocean south of Mauritius Island at approximately 152 hours . + +At 2 hours beyond pericynthion, a second descent propulsion maneuver was performed to shorten the return time and move the earth landing point to the South Pacific. The 263.8-second maneuver produced a velocity change of 860.5 ft/sec and resulted in an initial predicted earth landing point in the Pacific Ocean at 142:53:00. The transearth trip time was thus reduced by about 9 hours . +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +The analysis of the trajectory from lift-off to spacecraft/S-IVB separation was based on launch vehicle onboard data, as reported in reference 2, and from network tracking data. After separation, the actual trajectory information was determined from the best estimated trajectory generated from tracking and telemetry data. The earth and moon models used for the trajectory analysis are geometrically similar to those used for Apollo l2. Table 3-I is a listing of major flight events, and table 4-I defines the trajectory and maneuver parameters listed in table 4-II. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 19:56:20 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe accuracy of the first midcourse correction (table 4-Il), which placed the...', 'Result 1:\ncrewmen noted the small change in acceleration caused by the mixture ratio sh...', 'Result 1:\nThe accuracy of the first midcourse correction (table 4-Il), which placed the...', 'Result 1:\nThe planned launch and earth parking orbit phases for this mission were very ...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe planned launch and earth parking orbit phases for this mission were very ...', 'Result 1:\ninjection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866...', 'Result 1:\nThe analysis of the trajectory from lift-off to spacecraft/S-IVB separation w...', 'Result 1:\nThe analysis of the trajectory from lift-off to spacecraft/S-IVB separation w...'] +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:56:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.200, Perfect scores: 0/6 +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.83 +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:56:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:56:20 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.314, Max reward: 0.722 +2025-04-03 at 19:56:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:56:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What system was used to remove carbon dioxide from the lunar module during the Apollo missions? +2025-04-03 at 19:56:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 19:56:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo lunar module carbon dioxide scrubbing system" +2025-04-03 at 19:56:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:56:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the system used by the crew to remove carbon dioxide in the lunar module aboard the Apollo missions +2025-04-03 at 19:56:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:56:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:56:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar module ECS carbon dioxide removal system Apollo +2025-04-03 at 19:56:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 19:56:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what system used by apollo crew to remove carbon dioxide in lunar module compared to Sabatier process +2025-04-03 at 19:56:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:56:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:56:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How does the command module lithium hydroxide cartridge work? +2025-04-03 at 19:56:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The installed primary lithium hydroxide cartridge was used for approximately 27 hours (82 man-hoursīŧ‰ following activation of the lunar moduie at about 58 hours. The secondary cartridge was selected at about $\upbeta5$ 1/2 hours. During operation of the lunar module carbon dioxide removal system, the level was permitted to increase to an indicated l4.9 mm Hg. The primary cartridge is nominally rated for a usage capacity of 4l manhours at 520 Btu/man-hour. The secondary cartridge, nominally rated for about 17.9 man-hours, was used for 8 1/2 hours (25 1/2 man-hours). This cartridge is identical to that used in the portable life support system. A second primary cartridge was installed and used for approximately 6 minutes, but for the remainder of the mission, command module lithium hyaroxide cartridges were operated in a special arrangement. One side of each of two command module cartridges was covered and sealed with a plastic bag normally used to store a liquid-cooling garment. As shown +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-03 at 19:56:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo lunar module carbon dioxide removal mechanism +2025-04-03 at 19:56:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:56:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:56:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo command module lithium hydroxide cartridge system operation +2025-04-03 at 19:56:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The installed primary lithium hydroxide cartridge was used for approximately 27 hours (82 man-hoursīŧ‰ following activation of the lunar moduie at about 58 hours. The secondary cartridge was selected at about $\upbeta5$ 1/2 hours. During operation of the lunar module carbon dioxide removal system, the level was permitted to increase to an indicated l4.9 mm Hg. The primary cartridge is nominally rated for a usage capacity of 4l manhours at 520 Btu/man-hour. The secondary cartridge, nominally rated for about 17.9 man-hours, was used for 8 1/2 hours (25 1/2 man-hours). This cartridge is identical to that used in the portable life support system. A second primary cartridge was installed and used for approximately 6 minutes, but for the remainder of the mission, command module lithium hyaroxide cartridges were operated in a special arrangement. One side of each of two command module cartridges was covered and sealed with a plastic bag normally used to store a liquid-cooling garment. As shown +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-03 at 19:56:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:56:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo command module lithium hydroxide cartridge operation limitations +2025-04-03 at 19:56:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ +Result 2: +The installed primary lithium hydroxide cartridge was used for approximately 27 hours (82 man-hoursīŧ‰ following activation of the lunar moduie at about 58 hours. The secondary cartridge was selected at about $\upbeta5$ 1/2 hours. During operation of the lunar module carbon dioxide removal system, the level was permitted to increase to an indicated l4.9 mm Hg. The primary cartridge is nominally rated for a usage capacity of 4l manhours at 520 Btu/man-hour. The secondary cartridge, nominally rated for about 17.9 man-hours, was used for 8 1/2 hours (25 1/2 man-hours). This cartridge is identical to that used in the portable life support system. A second primary cartridge was installed and used for approximately 6 minutes, but for the remainder of the mission, command module lithium hyaroxide cartridges were operated in a special arrangement. One side of each of two command module cartridges was covered and sealed with a plastic bag normally used to store a liquid-cooling garment. As shown +------ + +2025-04-03 at 19:56:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:56:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lithium hydroxide cartridge system used to remove carbon dioxide during lunar module evacuation +2025-04-03 at 19:56:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 19:56:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:56:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How was the power system used to remove carbon dioxide using lithium hydroxide in the lunar module? +2025-04-03 at 19:56:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-03 at 19:56:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:56:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Reaction control system power requirements +2025-04-03 at 19:56:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:56:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:56:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Power system performance Apollo lunar module +2025-04-03 at 19:56:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:56:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:56:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module power system capacity +2025-04-03 at 19:56:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:56:41 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:56:41 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:56:41 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, True] +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1983, 400, 530, 342, 395, 895] +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [28, 28, 28, 28, 28, 28] +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_correctness:84 - Average student length: 757.50 +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 28.00 +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_correctness:86 - Length ratio: 27.05 +2025-04-03 at 19:56:41 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:56:41 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.325 Âą 0.368 +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.33 Âą 3.59 +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [10, 0, 1, 0, 0, 3] +2025-04-03 at 19:56:41 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The installed primary lithium hydroxide cartridge was used for approximately 27 hours (82 man-hoursīŧ‰ following activation of the lunar moduie at about 58 hours. The secondary cartridge was selected at about $\upbeta5$ 1/2 hours. During operation of the lunar module carbon dioxide removal system, the level was permitted to increase to an indicated l4.9 mm Hg. The primary cartridge is nominally rated for a usage capacity of 4l manhours at 520 Btu/man-hour. The secondary cartridge, nominally rated for about 17.9 man-hours, was used for 8 1/2 hours (25 1/2 man-hours). This cartridge is identical to that used in the portable life support system. A second primary cartridge was installed and used for approximately 6 minutes, but for the remainder of the mission, command module lithium hyaroxide cartridges were operated in a special arrangement. One side of each of two command module cartridges was covered and sealed with a plastic bag normally used to store a liquid-cooling garment. As shown +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The installed primary lithium hydroxide cartridge was used for approximately 27 hours (82 man-hoursīŧ‰ following activation of the lunar moduie at about 58 hours. The secondary cartridge was selected at about $\upbeta5$ 1/2 hours. During operation of the lunar module carbon dioxide removal system, the level was permitted to increase to an indicated l4.9 mm Hg. The primary cartridge is nominally rated for a usage capacity of 4l manhours at 520 Btu/man-hour. The secondary cartridge, nominally rated for about 17.9 man-hours, was used for 8 1/2 hours (25 1/2 man-hours). This cartridge is identical to that used in the portable life support system. A second primary cartridge was installed and used for approximately 6 minutes, but for the remainder of the mission, command module lithium hyaroxide cartridges were operated in a special arrangement. One side of each of two command module cartridges was covered and sealed with a plastic bag normally used to store a liquid-cooling garment. As shown +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ +Result 2: +The installed primary lithium hydroxide cartridge was used for approximately 27 hours (82 man-hoursīŧ‰ following activation of the lunar moduie at about 58 hours. The secondary cartridge was selected at about $\upbeta5$ 1/2 hours. During operation of the lunar module carbon dioxide removal system, the level was permitted to increase to an indicated l4.9 mm Hg. The primary cartridge is nominally rated for a usage capacity of 4l manhours at 520 Btu/man-hour. The secondary cartridge, nominally rated for about 17.9 man-hours, was used for 8 1/2 hours (25 1/2 man-hours). This cartridge is identical to that used in the portable life support system. A second primary cartridge was installed and used for approximately 6 minutes, but for the remainder of the mission, command module lithium hyaroxide cartridges were operated in a special arrangement. One side of each of two command module cartridges was covered and sealed with a plastic bag normally used to store a liquid-cooling garment. As shown +------ + +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ + +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds. + +About 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open. + +During passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\circ}$ to97°F。 + +6.6 DESCENT PROPULSION +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 19:56:41 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 19:56:41 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 19:56:41 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 19:56:41 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:56:41 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:56:41 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.220, Max reward: 0.590 +2025-04-03 at 19:56:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:56:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "S-IC boost maximum lateral load percentage" +2025-04-03 at 19:56:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 19:56:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"Studies on S-IC boost lateral load percentage in electric vehicles" +2025-04-03 at 19:56:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 19:56:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what percentage of maximum lateral loads experienced during S-IC boost compared to design value +2025-04-03 at 19:56:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-03 at 19:56:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ev S-IC boost lateral load percentage design value +2025-04-03 at 19:56:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-03 at 19:56:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:56:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: īŧ… of maximum lateral load experienced during S-IC boost +2025-04-03 at 19:56:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 19:56:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:56:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IC boost maximum lateral load experienced +2025-04-03 at 19:56:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 19:56:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:56:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: percentage of maximum lateral load experienced during S-IC boost compared to design value +2025-04-03 at 19:56:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-03 at 19:56:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:56:52 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:56:52 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:56:52 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, False, True, True] +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_correctness:82 - Student lengths: [433, 246, 492, 427, 109, 115] +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_correctness:84 - Average student length: 303.67 +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_correctness:86 - Length ratio: 30.37 +2025-04-03 at 19:56:52 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 19:56:52 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.308 Âą 0.270 +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.17 Âą 1.34 +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [1, 0, 1, 0, 4, 1] +2025-04-03 at 19:56:52 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 19:56:52 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 19:56:52 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +boost phases were within acceptable spacecraft structural. design limits. +------ + +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:56:52 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.33 +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:56:52 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:56:52 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.186, Max reward: 0.518 +2025-04-03 at 19:56:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:56:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "in-flight meal options airline selection criteria" +2025-04-03 at 19:56:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-03 at 19:56:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the basis for selecting flight menus +2025-04-03 at 19:56:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 19:56:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: history of airline menus UK BOAC +2025-04-03 at 19:56:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-03 at 19:56:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the factors that influence airline passengers when selecting flight options? +2025-04-03 at 19:56:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +TABLE 4-II.- TRAJECTORY PARAMETERS +------ +Result 2: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ + +2025-04-03 at 19:56:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "basis for selecting flight menus" +2025-04-03 at 19:56:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 19:56:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:57:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo mission 6 food selection process" +2025-04-03 at 19:57:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-03 at 19:57:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the principle behind selecting flight menus +2025-04-03 at 19:57:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 19:57:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo crew menu system rationing water +2025-04-03 at 19:57:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 19:57:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What criteria are airlines or travel websites typically use when developing flight menus? +2025-04-03 at 19:57:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 19:57:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "selection criteria for flight menus for spacecraft navigation" +2025-04-03 at 19:57:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 19:57:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:57:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 6 crew menu selection process" +2025-04-03 at 19:57:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:57:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Spacecraft attitude control reference +2025-04-03 at 19:57:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-03 at 19:57:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Water conservation strategies during lunar missions +2025-04-03 at 19:57:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 19:57:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo crew meal preferences and logistics +2025-04-03 at 19:57:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 19:57:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "basis for selecting flight menus for spacecraft attitude control" +2025-04-03 at 19:57:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 19:57:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:57:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no search needed +2025-04-03 at 19:57:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:57:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: attitude control and platform alignment in spacecraft +2025-04-03 at 19:57:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-03 at 19:57:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Chlorination and bacterial growth during space missions +2025-04-03 at 19:57:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 19:57:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "attitude control cockpit menu options astronaut spaceflight" +2025-04-03 at 19:57:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 19:57:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:57:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 6 crew menu planning and selection process" +2025-04-03 at 19:57:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-03 at 19:57:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how do crew select flight menus +2025-04-03 at 19:57:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 19:57:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar module water usage rates +2025-04-03 at 19:57:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-03 at 19:57:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar mission attitude control manual alignment" +2025-04-03 at 19:57:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-03 at 19:57:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:57:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no search needed +2025-04-03 at 19:57:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:57:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: meals vs menu system for astronauts +2025-04-03 at 19:57:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 19:57:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Energy consumption patterns during space missions +2025-04-03 at 19:57:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:57:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "manuale propulsion maneuver lunar module attitude control astronaut training" +2025-04-03 at 19:57:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 19:57:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:57:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no search needed +2025-04-03 at 19:57:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:57:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "primary guidance attitude control maneuvers" +2025-04-03 at 19:57:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 19:57:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:57:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no search needed +2025-04-03 at 19:57:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:57:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "primary guidance system performance apollo spacecraft" +2025-04-03 at 19:57:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:57:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:57:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no search needed +2025-04-03 at 19:57:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:57:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo spacecraft flight menu selection criteria" +2025-04-03 at 19:57:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:57:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:57:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no search needed +2025-04-03 at 19:57:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:57:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo astronaut crew operations manual" +2025-04-03 at 19:57:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:57:24 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:57:24 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:57:24 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, False, True, False] +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_correctness:82 - Student lengths: [372, 432, 485, 558, 612, 2033] +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [82, 82, 82, 82, 82, 82] +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_correctness:84 - Average student length: 748.67 +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 82.00 +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_correctness:86 - Length ratio: 9.13 +2025-04-03 at 19:57:24 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-03 at 19:57:24 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.662 Âą 0.347 +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 9.50 Âą 9.01 +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [10, 0, 6, 28, 3, 10] +2025-04-03 at 19:57:24 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-03 at 19:57:24 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ +Result 2: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident. + +7.2.l Des cent Propulsion Propellants + +The loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off. +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +TABLE 4-II.- TRAJECTORY PARAMETERS +------ +Result 2: +The extensive testing and analyses and the consistency with which the postflight test data repeated preflight acceptance test results indicate the problem was most likely caused either by the Command Module Pilot responding_too quickly to the $0.05\tt{e}$ light not coming on or by an intermittent hardware failure that cleared itself during entry. + +Based on these findings, a change is not warranted to existing procedures or hardware on future flights. + +This anomaly is closed. + +14.1.6 Gas Leak in Apex Cover Jettison System +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +------ +Result 2: +The crew followed the flight menus prior to the inflight incident and maintained a complete log of foods consumed.¡ To conserve water during the abort phase, the crew consumed only those foods which did not require water for rehydration. The crew drank juices in preference to plain water to help maintain their electrolyte balance. + +The crew's comments about the quality of the food were generally favorable, but they reported that food packaging and stowage could be improved. The crew encountered some difficulty in removing the meal packages from the lower equipment bay food container and in replacing Some uneaten food items. Preflight briefings of future crews should alleviate these difficulties. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress. + +9.2.5 Food + +The flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches. +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +to command rotation about the vehicle pitch and roll axes and the attitude controller for yaw commands. The pilot's task was further complicated by having the flight director attitude indicators powered down. Without these displays, it was necessary to monitor attitudes by observing gimbal angles on the display and keyboard assembly. Because the spacecraft yaw axis was not coincident to that of the platform yaw axis, either a pitch or roll command would cause a change in both of the corresponding gimbal-angle displays. After the vehicle attitude was changed to more closely align with the platform and to reduce the yaw gimbalangle disparity, passive thermal control was established satisfactorily. Both guidance systems were then powered down until l05 hours. At that time, the abort guidance system was powered up for control during the first transearth midcourse correction. The passive thermal control mode was reestablished and the abort system was powered down. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +To conserve reaction control fuel when holding an attitude, a wide deadband was established using primary guidance. Because the platform was not aligned with & passive thermal control mode reference matrix, yawing the vehicle each hour resulted in inner and middle gimbal angle deviations. The crew could not determine any standard procedure to keep the middle angle constant during the maneuver. As the spacecraft maneuvered from one quadrant to the next, the same thrust/translation controller assembly input wouid result in a different effect in controlling the middle gimb al angle. + +8.7.5 Platform Alignment +------ +Result 2: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +started to sight the service module in the docking window. The lightened spacecraft combination was easily maneuvered using attitude control in both the manual minimum-impulse and automatic attitude-hold modes. +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +Initial outside observations through the lunar module windows indicated that normal. platform aligmments using a star reference would be extremely difficult because of the large amoumt of debris in the vicinity of the spacecraft. This debris apparently originated during the tank incident. A subsequent observation when the spacecraft was in the moon's shadow indicated that an alignment at that time would have been feasible because of the improved visual contrast. Crew training for sun/earth and Sum/moon alignments in the simulators should be emphasized to handle contingencies such as occurred during Apollo l3. + +8.8 TRANSEARTH INJECTION + +Maneuvering to the proper attitude for transearth injection was done manually with the thrust/translation controller assembly and attitude controller assembly while tracking primary guidance error needles. The error needles were nulled, and the spacecraft was then placed in the primary guidance automatic control mode to maintain attitude. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:57:24 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nto command rotation about the vehicle pitch and roll axes and the attitude co...', 'Result 1:\nPlatform-sensed velocity changes, ft/sec Command module axes Lunar module axe...', 'Result 1:\nTo conserve reaction control fuel when holding an attitude, a wide deadband w...', 'Result 1:\nstarted to sight the service module in the docking window. The lightened spac...', 'Result 1:\nAt approximately 105 hours, the crew performed a manual descent propulsion ma...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nA descent propulsion system maneuver to reestablish a free-return trajectory ...', 'Result 1:\nThe vehicle was launched on an azimuth 90 degrees east of north, and a roll m...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...'] +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:57:24 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.233, Perfect scores: 0/6 +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.83 +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:57:24 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:57:24 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.331, Max reward: 0.670 +2025-04-03 at 19:57:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:57:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: shaft rotation zero output equivalent +2025-04-03 at 19:57:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 19:57:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: equivalent of zero rotation shaft output +2025-04-03 at 19:57:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 19:57:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "standstill engine rotation equivalent +2025-04-03 at 19:57:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 19:57:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the equivalent of zero output in shaft rotation +2025-04-03 at 19:57:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 19:57:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:57:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: zero optics mode nulling circuit +2025-04-03 at 19:57:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 19:57:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: difference between half-speed resolver and optics shaft servo loop mechanization +2025-04-03 at 19:57:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 19:57:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:57:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: zero optics mode motor drive amplifier circuitry +2025-04-03 at 19:57:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-03 at 19:57:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: difference between half-speed resolver and normal operation on shaft axis +2025-04-03 at 19:57:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 19:57:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:57:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: half-speed resolver shaft axis definition +2025-04-03 at 19:57:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 19:57:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:57:39 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:57:39 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:57:39 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, True, False, True, False, False] +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_correctness:82 - Student lengths: [697, 427, 458, 648, 536, 700] +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [12, 12, 12, 12, 12, 12] +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_correctness:84 - Average student length: 577.67 +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 12.00 +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_correctness:86 - Length ratio: 48.14 +2025-04-03 at 19:57:39 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 19:57:39 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.358 Âą 0.299 +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.50 Âą 1.50 +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [1, 3, 1, 4, 0, 0] +2025-04-03 at 19:57:39 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 19:57:39 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not ...'] +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 19:57:39 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 19:57:39 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 19:57:39 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.167, Perfect scores: 1/6 +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.50 +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 19:57:39 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:57:39 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.291, Max reward: 0.709 +2025-04-03 at 19:57:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:57:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: publication date of performance analysis report by notable companies 2022 +2025-04-03 at 19:57:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 19:57:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: report titled "Performance Analysis" publication date +2025-04-03 at 19:57:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 19:57:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: publication date "Performance Analysis report" +2025-04-03 at 19:57:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 19:57:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: publication date report "Performance Analysis" +2025-04-03 at 19:57:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 19:57:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:57:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission report performance analysis publication date +2025-04-03 at 19:57:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:57:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission reports Performance Analysis publication date +2025-04-03 at 19:57:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:57:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 7-13 Performance Analysis report publication date +2025-04-03 at 19:57:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:57:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: performance of lunar module reaction control system final review +2025-04-03 at 19:57:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:57:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:57:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission report supplement performance analysis publication date +2025-04-03 at 19:57:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 19:57:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission Performance Analysis publication date +2025-04-03 at 19:57:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:57:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system final review 1970 +2025-04-03 at 19:57:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:57:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:57:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 Performance Analysis supplement publication date +2025-04-03 at 19:57:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:57:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission Performance Analysis report not found +2025-04-03 at 19:57:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 19:57:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:57:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission report Performance Analysis publication December 1969 +2025-04-03 at 19:57:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:57:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 7 supplemental report with Performance Analysis +2025-04-03 at 19:57:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:57:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:57:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 7 report title match +2025-04-03 at 19:57:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:57:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:57:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 9 and 10 Command and Service Module reviews +2025-04-03 at 19:57:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:57:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:57:55 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:57:55 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:57:55 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, True, True] +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_correctness:82 - Student lengths: [340, 503, 395, 622, 246, 331] +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [13, 13, 13, 13, 13, 13] +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_correctness:84 - Average student length: 406.17 +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 13.00 +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_correctness:86 - Length ratio: 31.24 +2025-04-03 at 19:57:55 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 19:57:55 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.508 Âą 0.393 +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.83 Âą 2.54 +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [5, 2, 0, 0, 7, 3] +2025-04-03 at 19:57:55 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:57:55 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...'] +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-03 at 19:57:55 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-03 at 19:57:55 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ + +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ + +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:57:55 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nand Service Module Reaction Control System Apri1 1970 5 Service Propulsion Sy...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...'] +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:57:55 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.167, Perfect scores: 0/6 +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.00 +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:57:55 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:57:55 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.370, Max reward: 0.706 +2025-04-03 at 19:57:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:57:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: customers to search if the above query did not return results, I broaden my search to: "NASAdocument approval year 1970s-2000s" +2025-04-03 at 19:57:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:57:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Artemis Mission Evaluation Team approved document year" +2025-04-03 at 19:57:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:57:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What year was the Space Shuttle Mission Evaluation Report approved by the NASA Mission Evaluation Team? +2025-04-03 at 19:57:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 19:57:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: mission evaluation team document approval year +2025-04-03 at 19:57:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 19:57:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:58:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Artemis Mission Evaluation Team approved document September 1970" +2025-04-03 at 19:58:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:58:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Mission Report approval date +2025-04-03 at 19:58:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 19:58:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:58:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 preliminary document approved by Mission Evaluation Team September 1970" +2025-04-03 at 19:58:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:58:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Abraham Waldron Apollo 13 Mission Report approval +2025-04-03 at 19:58:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:58:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:58:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 training programã“ãĄã‚‰ evaluation report approved by Mission Evaluation Team September 1970" +2025-04-03 at 19:58:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:58:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Abraham Waldron Apollo 13 mission report +2025-04-03 at 19:58:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:58:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:58:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 report approved by Mission Evaluation Team 1970" +2025-04-03 at 19:58:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:58:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Abraham Waldron mission evaluation report Apollo 13 +2025-04-03 at 19:58:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:58:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:58:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 document approved by Mission Evaluation Team September 1970" +2025-04-03 at 19:58:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:58:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: John C. Stevens and Walter R. Weems Apollo 13 mission report +2025-04-03 at 19:58:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:58:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:58:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 12 document approved by Mission Evaluation Team September 1970" +2025-04-03 at 19:58:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:58:14 | INFO | src.agent:run_tool_calls:115 - īŋŊīŋŊīŋŊ Search Query: Abraham Waldron Apollo 7 mission report +2025-04-03 at 19:58:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:58:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:58:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 12 document approved by Mission Evaluation Team" +2025-04-03 at 19:58:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:58:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: hard copy Apollo 7 Mission Manual +2025-04-03 at 19:58:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:58:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:58:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Wernher von Braun Apollo 11 Mission Report +2025-04-03 at 19:58:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:58:18 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:58:18 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:58:19 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, True, False, False] +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_correctness:82 - Student lengths: [28, 436, 205, 400, 103, 2034] +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_correctness:84 - Average student length: 534.33 +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 4.00 +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_correctness:86 - Length ratio: 133.58 +2025-04-03 at 19:58:19 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:58:19 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.417 Âą 0.398 +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 3.33 Âą 3.73 +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 2, 0, 8, 1, 9] +2025-04-03 at 19:58:19 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 19:58:19 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 19:58:19 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 19:58:19 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.067, Perfect scores: 0/6 +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:58:19 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:58:19 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.238, Max reward: 0.950 +2025-04-03 at 19:58:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:58:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA manned space center location +2025-04-03 at 19:58:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Johnson Space Center location +2025-04-03 at 19:58:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 19:58:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Lyndon B. Johnson Space Center location +2025-04-03 at 19:58:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:58:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "John F. Kennedy Space Center location +2025-04-03 at 19:58:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "nasa manned space center location +2025-04-03 at 19:58:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:58:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: postal address of NASA manned space center +2025-04-03 at 19:58:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: address of NASA Manned Spacecraft Center Houston +2025-04-03 at 19:58:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lyndon B. Johnson Space Center map +2025-04-03 at 19:58:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:58:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Manned Spacecraft Center history +2025-04-03 at 19:58:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Manned Spacecraft Center Houston Texas 1970 +2025-04-03 at 19:58:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nearby locations to Lyndon B. Johnson Space Center +2025-04-03 at 19:58:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:58:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Space Center Houston +2025-04-03 at 19:58:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:58:32 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:58:32 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:58:32 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, False, False, False] +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_correctness:82 - Student lengths: [333, 439, 427, 147, 371, 120] +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [20, 20, 20, 20, 20, 20] +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_correctness:84 - Average student length: 306.17 +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 20.00 +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_correctness:86 - Length ratio: 15.31 +2025-04-03 at 19:58:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_format:228 - Responses ending properly: 1/6 +2025-04-03 at 19:58:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.425 Âą 0.253 +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.17 Âą 1.57 +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [3, 4, 0, 4, 1, 1] +2025-04-03 at 19:58:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-03 at 19:58:32 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 5.0 +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.833 +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:58:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.033, Perfect scores: 0/6 +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:58:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:58:32 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.300, Max reward: 0.535 +2025-04-03 at 19:58:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:58:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot function description NASA space mission +2025-04-03 at 19:58:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +The Apollo l3 space vehicle was launched from pad A of complex 39, Kennedy Space Center, Florida. Except for the high-amplitude, lowfrequency oscillations which resulted in premature cutoff of the S-II center engine, the basic performance of the launch vehicle was normal. Despite the anomaly, all launch vehicle objectives were achieved, as discussed in reference 2. In addition, the S-IvB lunar impact experiment was accomplished, as discussed in section ll.3. +------ + +2025-04-03 at 19:58:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot mission function description +2025-04-03 at 19:58:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:58:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Was pilot involved in describing function during mission Apollo +2025-04-03 at 19:58:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ + +2025-04-03 at 19:58:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Was pilot involved in describing a function during space mission? describe pilot involvement during spacewalk or EVA +2025-04-03 at 19:58:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ + +2025-04-03 at 19:58:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot mission function description +2025-04-03 at 19:58:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:58:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:58:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 mission report NASA +2025-04-03 at 19:58:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:58:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission pilot roles and functions during mission +2025-04-03 at 19:58:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:58:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: descritporsch pilot and crew reaction during cryogenic oxygen tank incident +2025-04-03 at 19:58:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal. + +During the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l. +------ + +2025-04-03 at 19:58:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: + + +2025-04-03 at 19:58:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:58:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:58:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 mission report NASA +2025-04-03 at 19:58:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:58:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission post-flight test procedures +2025-04-03 at 19:58:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:58:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Describe critical decision-making by lunar module pilot during cryogenic oxygen tank failure +2025-04-03 at 19:58:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 19:58:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:58:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Neil Armstrong Apollo 11 +2025-04-03 at 19:58:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:58:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module pilot role during post-flight +2025-04-03 at 19:58:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:58:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Describes lunar module pilot's role in aborting mission and managing crisis situation +2025-04-03 at 19:58:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 19:58:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:58:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: James McDivitt Apollo 9 +2025-04-03 at 19:58:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ + +2025-04-03 at 19:58:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module pilot description of mission phases +2025-04-03 at 19:58:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:58:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module and command module pilot during critical abort sequence +2025-04-03 at 19:58:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 19:58:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:58:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 similarity to Apollo 11 +2025-04-03 at 19:58:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:58:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module pilot role in mission phases +2025-04-03 at 19:58:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:58:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module undocking chalides and navigation challenges +2025-04-03 at 19:58:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:58:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:58:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 pilot function description +2025-04-03 at 19:58:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:58:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module pilot procedures for lunar landing +2025-04-03 at 19:58:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:58:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module undocking pilot coordination and spacecraft navigation during entry +2025-04-03 at 19:58:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:58:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 19:58:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Neil Armstrong Edwin Buzz Aldrin Apollo 11 +2025-04-03 at 19:58:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:58:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar module pilot role during lunar landing +2025-04-03 at 19:58:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:58:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module undocking and entry procedures emphasis on safety and systems management +2025-04-03 at 19:58:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:58:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:58:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 crew roles during lunar landing +2025-04-03 at 19:58:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:58:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module pilot expertise in undocking and entry phase +2025-04-03 at 19:58:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:58:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:58:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 lunar landing pilot role +2025-04-03 at 19:58:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:58:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module pilot resource management during critical phases +2025-04-03 at 19:58:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 19:58:59 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:58:59 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:58:59 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_correctness:82 - Student lengths: [360, 1730, 1739, 174, 1633, 309] +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [3, 3, 3, 3, 3, 3] +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_correctness:84 - Average student length: 990.83 +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 3.00 +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_correctness:86 - Length ratio: 330.28 +2025-04-03 at 19:58:59 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 19:58:59 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.617 Âą 0.365 +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 5.17 Âą 4.26 +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [1, 8, 10, 0, 10, 2] +2025-04-03 at 19:58:59 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted. + +The Apollo l3 space vehicle was launched from pad A of complex 39, Kennedy Space Center, Florida. Except for the high-amplitude, lowfrequency oscillations which resulted in premature cutoff of the S-II center engine, the basic performance of the launch vehicle was normal. Despite the anomaly, all launch vehicle objectives were achieved, as discussed in reference 2. In addition, the S-IvB lunar impact experiment was accomplished, as discussed in section ll.3. +------ + +2025-04-03 at 19:58:59 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...'] +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 19:58:59 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\n1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJEC...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nMedical kits for future flights will include nose drops packaged the same as ...'] +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 19:58:59 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\n1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJEC...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...'] +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-03 at 19:58:59 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal. + +During the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l. +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions: + +a. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The maneuver to the undocking attitude was made by the lunar module. Time consuming operations were followed to avoid gimbal lock of both spacecraft platforms. Because of the difference in alignments between the two spacecraft, considerable difficul.ty was encountered in maneuvering to the lunar module undocking attitude without driving the command module platform into gimbal lock. The maneuver required a complicated procedure using the lunar module platform and close cooperation between the Commander and Command Module Pilot. The resulting maneuver also used up considerable lunar module reaction control fuel. The final undocking attitude was very close to command module gimbal lock attitude. A different command module alignment procedure should have been used to prevent the probability of gimbal lock. +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 19:58:59 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\ndiscussion of the flight and associated crew activities is presented in secti...', 'Result 1:\nThe two tanks cortaining cryogenic oxygen, used for _fuel cell operation and ...', 'Result 1:\nThe two tanks cortaining cryogenic oxygen, used for _fuel cell operation and ...', 'Result 1:\na. The mission was aborted because of the total loss of primary oxygen in the...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...', 'Result 1:\nThe maneuver to the undocking attitude was made by the lunar module. Time con...', 'Result 1:\nA major flight control function, in adaition to the monitoring of systems sta...', 'Result 1:\nA major flight control function, in adaition to the monitoring of systems sta...', 'Result 1:\nA major flight control function, in adaition to the monitoring of systems sta...'] +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 19:58:59 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\n1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJEC...', 'Result 1:\n(section ll.3).\n------\nResult 2:\n+21.0 +3.0\n------\n...'] +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 19:58:59 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.233, Perfect scores: 0/6 +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 2.33 +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:58:59 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:58:59 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.398, Max reward: 0.800 +2025-04-03 at 19:59:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:59:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How much heat dissipated by 50 pounds of water when reduced to background power consumption conditions +2025-04-03 at 19:59:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"thermal load and cooling capacity cooling systems" +2025-04-03 at 19:59:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:59:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +This calculation can be done using the formula for melting rate, considering initial water temperature to find its final temperature after flow results in reduced cooling. +2025-04-03 at 19:59:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-03 at 19:59:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: cooling water for liquids usually meaningless" +2025-04-03 at 19:59:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "50 pound water mass cooling capacity hours +2025-04-03 at 19:59:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 19:59:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:59:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: thermal management and cooling with potable water at 50 pounds +2025-04-03 at 19:59:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"lithium-ion battery thermal management systems" +2025-04-03 at 19:59:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 19:59:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What is the normal operating level or range for a potable water tank in a spacecraft? +2025-04-03 at 19:59:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "source melting point-dissolution coefficients- scale-on hot condensate" +2025-04-03 at 19:59:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 19:59:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 40-psi water cooling system cooling capacity hours +2025-04-03 at 19:59:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 19:59:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:59:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: potable water freezing and insulation effect on crew comfort at 50 pounds +2025-04-03 at 19:59:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Potable water was obtained by periodically pressurizing the potable tank with surge-tank oxygen and withdrawing potable water until the pressures equalized. This method provided potable water for crew use until 24 hours prior to entry, at which time water could not be withdrawn from the potable tank and it appeared to be exhausted [section 5.8]. + +The hatch, probe, and drogue were secured in the couches by lap belt and shoulder harness restraints to prevent movement during subsequent maneuvers. + +8.7.4 Midcourse Correction to a Free Return +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"lithium-ion battery cooling system lunar mission" +2025-04-03 at 19:59:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 19:59:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What was the temperature at which water could no longer be withdrawn from the potable tank after 120 hours in space? +2025-04-03 at 19:59:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 19:59:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Vacuum systems purge procedures" +2025-04-03 at 19:59:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:59:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 20-psi system cooling capacity 50 pound water +2025-04-03 at 19:59:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:59:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:59:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effect of insulation and thermal management on cooling a 50-pound tank of water +2025-04-03 at 19:59:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 19:59:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"lithium-ion battery thermal management system in space missions" +2025-04-03 at 19:59:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 19:59:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What is the effect of water freezing in the lines of a spacecraft potable water system on the operation of the system? +2025-04-03 at 19:59:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (health and performance and feedback (water, parts ) +2025-04-03 at 19:59:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-03 at 19:59:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 40-psi system water cooling time 50 pounds +2025-04-03 at 19:59:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 19:59:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:59:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fusion heating and contamination of cryogenic tank 2 +2025-04-03 at 19:59:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 19:59:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"lithium-ion battery thermal management in Apollo missions" +2025-04-03 at 19:59:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:59:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What is the temperature below which the command module structure near the tank and lines is cold to touch? +2025-04-03 at 19:59:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ + +2025-04-03 at 19:59:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Lunar module water capabilities comparison +2025-04-03 at 19:59:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:59:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 22 minutes water cooling capability 50 pounds +2025-04-03 at 19:59:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:59:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:59:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: potential for electrical overheating in cryogenic tank filling operations +2025-04-03 at 19:59:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 19:59:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"lithium-ion power consumption thermal management" +2025-04-03 at 19:59:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 19:59:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What is the temperature range of acceptable low injector temperature for engine operation? +2025-04-03 at 19:59:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ + +2025-04-03 at 19:59:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "mechanisms for preventing water intrusion under command module operation" +2025-04-03 at 19:59:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the periods when it was activated, the command module environmental control system performed normally. From the time of powering dowm at approximately 58 hours until reactivation approximately l-1/2 hours before entry, environmental control for the interconnected cabins was maintained using lumar module equipment. Two anomalies associated with the environmental control instrumentation occurred and are discussed in sections 14.l.8 and l4.l.9. An additional discrepancy, noted after landing and discussed in section l0.3, was the position of the inlet postlanding ventilation valve at the time of recovery. This discrepancy is discussed in section 14.l.2. +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:59:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 50 pound water cooling time hours +2025-04-03 at 19:59:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Potable water was obtained by periodically pressurizing the potable tank with surge-tank oxygen and withdrawing potable water until the pressures equalized. This method provided potable water for crew use until 24 hours prior to entry, at which time water could not be withdrawn from the potable tank and it appeared to be exhausted [section 5.8]. + +The hatch, probe, and drogue were secured in the couches by lap belt and shoulder harness restraints to prevent movement during subsequent maneuvers. + +8.7.4 Midcourse Correction to a Free Return +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:59:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:59:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: improvement to tank insulation and electrical protection during liquid nitrogen and liquid oxygen storage +2025-04-03 at 19:59:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 19:59:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"lithium-ion battery cooling systems volumetric efficiency" +2025-04-03 at 19:59:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 19:59:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What is normal temperature range for engine package during peak activity? +2025-04-03 at 19:59:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:59:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Difference between lunar module and command module functions" +2025-04-03 at 19:59:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 19:59:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 24 hour water cooling capability 50 pounds +2025-04-03 at 19:59:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 19:59:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:59:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: oxygen consumption during spacecraft depressurization +2025-04-03 at 19:59:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 19:59:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"50 pounds of water as a cooling system" +2025-04-03 at 19:59:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 19:59:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What is the normal cabin temperature range during the return flight when power levels are maintained between 350-400 watts? +2025-04-03 at 19:59:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "performance normality of systems due to Jupiter flyby effects" +2025-04-03 at 19:59:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +With the exception of supercritical helium system performance, descent propulsion system operation, including engine starts and throttle response, was normal. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 19:59:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 19:59:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effective power consumption of mechanical systems during spacewalk +2025-04-03 at 19:59:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:59:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What is the normal temperature range for the structure near the tank and lines? +2025-04-03 at 19:59:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:59:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 19:59:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fuel cell power consumption and usage during lunar module ascent and free return +2025-04-03 at 19:59:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:59:42 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 19:59:42 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 19:59:43 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 1/6 answers correct +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, True] +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_correctness:82 - Student lengths: [2004, 389, 1543, 1281, 1215, 365] +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_correctness:84 - Average student length: 1132.83 +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 8.00 +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_correctness:86 - Length ratio: 141.60 +2025-04-03 at 19:59:43 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 19:59:43 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.554 Âą 0.326 +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 25.67 Âą 36.32 +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 3/6 +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [10, 0, 15, 16, 106, 7] +2025-04-03 at 19:59:43 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Potable water was obtained by periodically pressurizing the potable tank with surge-tank oxygen and withdrawing potable water until the pressures equalized. This method provided potable water for crew use until 24 hours prior to entry, at which time water could not be withdrawn from the potable tank and it appeared to be exhausted [section 5.8]. + +The hatch, probe, and drogue were secured in the couches by lap belt and shoulder harness restraints to prevent movement during subsequent maneuvers. + +8.7.4 Midcourse Correction to a Free Return +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Total oxygen usage from the three lunar module oxygen tanks was 20.3 pounds over an 82-hour period, for an average usage rate of 0.25 lb/ hr. Oxygen remaining in the tanks at undocking was 33.5 pounds. During translunar coast, lunar module cabin leakage was about 0.0l4 lb/hr, assuming an average cabin pressure of 4.5 psia. Command module cabin leakage was estimated to have been about 0.027 lb/hr. These values indicate an average metabolic consumption rate throughout the flight of approximately 0.21 1b/hr. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained. + + + +Figure 7.l-l.- Entry battery energy. + +7.2 LUNAR MODULE + +Following lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of + +7.1.3 Cryogenic Fluids + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls。 +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:59:43 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nfreezing in the lines most likely occurred. This freezing condition could hav...', 'Result 1:\nPotable water was obtained by periodically pressurizing the potable tank with...', 'Result 1:\nThe most likely cause of the anomaly is a tank-insulation degradation which w...', 'Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...', 'Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...', 'Result 1:\nAfter initial cryogenic oxygen filling during the countdown demonstration tes...', 'Result 1:\nTotal oxygen usage from the three lunar module oxygen tanks was 20.3 pounds o...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nThe command module was completely powered down at 58 hours 40 minutes , at wh...'] +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-03 at 19:59:43 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ +Result 2: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The battery potting will be improved to prevent electrolyte bridging between the battery terminals and case. These improvements include coating the inside of the battery case with epoxy paint before the battery is assembled and changing the potting material used at the ends of the case to a material which has better adhesion characteristics. Also, the cell chimneys will be manifolded together and to the case vent-valve with plastic tubing. + +In addition, tests are being performed to determine if the quantity of free electrolyte in each cell can be reduced. Preliminary results indicate a reduction of from 360 to 340 cc per cell is possible. +------ +Result 2: +The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 19:59:43 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nThe electrical power system performed all required functions. At lunar module...', 'Result 1:\nAt the time of power up, 2l79 ampere-hours of electrical energy was available...', 'Result 1:\nadvised of their consumables status. A procedure was developed on the ground ...', 'Result 1:\nThe battery potting will be improved to prevent electrolyte bridging between ...', 'Result 1:\nfreezing in the lines most likely occurred. This freezing condition could hav...'] +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ +Result 2: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:59:43 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nfreezing in the lines most likely occurred. This freezing condition could hav...', 'Result 1:\n6.7 ENVIRONMENTAL CONTROL\n\nEnvironmental. control system performance was sati...', 'Result 1:\nDuring the period when the command module was powered down, the cabin tempera...', 'Result 1:\nfreezing in the lines most likely occurred. This freezing condition could hav...', 'Result 1:\nDuring the period when the command module was powered down, the cabin tempera...', 'Result 1:\nThe command module reaction control system helium pressures and temperatures ...', 'Result 1:\nDuring the peak engine activity period after the oxygen tank incident, engine...', 'Result 1:\nLow cabin temperature, resulting from a greatly reduced thermal loading from ...', 'Result 1:\nDuring the peak engine activity period after the oxygen tank incident, engine...'] +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +During the periods when it was activated, the command module environmental control system performed normally. From the time of powering dowm at approximately 58 hours until reactivation approximately l-1/2 hours before entry, environmental control for the interconnected cabins was maintained using lumar module equipment. Two anomalies associated with the environmental control instrumentation occurred and are discussed in sections 14.l.8 and l4.l.9. An additional discrepancy, noted after landing and discussed in section l0.3, was the position of the inlet postlanding ventilation valve at the time of recovery. This discrepancy is discussed in section 14.l.2. +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +A.2 LUNAR MODULE +------ +Result 2: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +With the exception of supercritical helium system performance, descent propulsion system operation, including engine starts and throttle response, was normal. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 19:59:43 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nfreezing in the lines most likely occurred. This freezing condition could hav...', 'Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...', 'Result 1:\nSystem decontamination at Hawaii was normal, except that the system l fuel is...', 'Result 1:\n6.7 ENVIRONMENTAL CONTROL\n\nEnvironmental. control system performance was sati...', 'Result 1:\n6.7 ENVIRONMENTAL CONTROL\n\nEnvironmental. control system performance was sati...', 'Result 1:\nDuring the periods when it was activated, the command module environmental co...', 'Result 1:\nA.2 LUNAR MODULE\n------\nResult 2:\nPlatform-sensed velocity changes, ft/sec Co...', 'Result 1:\nWith the exception of supercritical helium system performance, descent propul...'] +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4). + +7.2.4 Water + +During the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking¡is shown in figure 7.2-l. + +7.2.5 Batteries +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Potable water was obtained by periodically pressurizing the potable tank with surge-tank oxygen and withdrawing potable water until the pressures equalized. This method provided potable water for crew use until 24 hours prior to entry, at which time water could not be withdrawn from the potable tank and it appeared to be exhausted [section 5.8]. + +The hatch, probe, and drogue were secured in the couches by lap belt and shoulder harness restraints to prevent movement during subsequent maneuvers. + +8.7.4 Midcourse Correction to a Free Return +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 1.0 +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.167 +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 19:59:43 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.17 +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 19:59:43 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 19:59:43 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.348, Max reward: 0.826 +2025-04-03 at 19:59:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:59:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module shuttle hot water port nickel content +2025-04-03 at 19:59:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module hot water loop systems nickel content +2025-04-03 at 19:59:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "command module hot water port nickel content" +2025-04-03 at 19:59:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was nickel content in space shuttle Columbia's Command Module Toilet (CM Hot Water Port) +2025-04-03 at 19:59:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel content command module hot water port space station +2025-04-03 at 19:59:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel content command module hot water port +2025-04-03 at 19:59:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:59:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel content command module hot water port space shuttle +2025-04-03 at 19:59:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module hot water port nickel sample context Apollo 11 +2025-04-03 at 19:59:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:59:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo command module hot water port nickel content primary source +2025-04-03 at 19:59:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 19:59:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel content in Command Module Hot Water Port after 1970 Apollo mission +2025-04-03 at 19:59:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 19:59:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel content command module water supply Apollo missions +2025-04-03 at 19:59:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:59:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo command module hot water port nickel content +2025-04-03 at 19:59:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 19:59:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 19:59:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module hot water port nickel analysis experiment +2025-04-03 at 19:59:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ + +2025-04-03 at 19:59:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo command module hot water port nickel content +2025-04-03 at 19:59:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 19:59:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 nickel content in Command Module +2025-04-03 at 19:59:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 19:59:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel content surface material Saturn command module Apollo +2025-04-03 at 19:59:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:59:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo command module hot water port sample analysis +2025-04-03 at 19:59:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:59:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 19:59:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module hot water port system nickel analysis spacecraft +2025-04-03 at 19:59:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 19:59:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 command module reaction control system propellant, nickel content +2025-04-03 at 19:59:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 19:59:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel content spacesuit instrumentation material electroless plating +2025-04-03 at 19:59:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 19:59:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module hot water port nickel test results +2025-04-03 at 19:59:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 19:59:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module hot water port nickel guideline NASA +2025-04-03 at 19:59:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 19:59:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel content Apollo 13 Service Module hot water port +2025-04-03 at 19:59:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 19:59:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module insulation materials Apollo missions +2025-04-03 at 19:59:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 19:59:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:00:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module hot water port nickel plot engineering description +2025-04-03 at 20:00:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-03 at 20:00:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel content in lunar module hot water port +2025-04-03 at 20:00:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 20:00:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module hot water port nickel content Apollo mission +2025-04-03 at 20:00:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:00:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:00:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: dissimilar metal effect hot water port nickel +2025-04-03 at 20:00:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The transducer is a variable reluctance instrument actuated by differential pressure applied across a twisted Bourdon tube. The housing, including the cavity containing the Bourdon tube and the variable reluctance elements, is nickel plated. The Bourdon tube-variable reluctance + +assembly and the sense port fitting are soldered in place. Inspection of the failed units indicates that the flaking occurs adjacent to the solder. + +The most probable cause of the problem is poor plating adhesion to the aluminum base metal. Differential expansion between the solder and the aluminum may cause the plating to crack. Moisture from the environmental control system suit loop could then penetrate the plating, corrode the aluminum base metal and cause the plating to peel and flake. The nickel flakes could then enter the air gap of the variable reluctance elements and affect the measurement. +------ + +2025-04-03 at 20:00:04 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Command Module System failure +2025-04-03 at 20:00:04 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 20:00:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:00:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel plating failure investigation command module +2025-04-03 at 20:00:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:00:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel content in command module oxygen tank +2025-04-03 at 20:00:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 20:00:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:00:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel content in oxygen tank +2025-04-03 at 20:00:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ +Result 2: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ + +2025-04-03 at 20:00:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:00:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nickel content Apollo 13 command module +2025-04-03 at 20:00:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 20:00:10 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:00:10 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:00:10 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, True, False, True, True] +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.67 +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_correctness:82 - Student lengths: [571, 353, 302, 1210, 361, 443] +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [20, 20, 20, 20, 20, 20] +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_correctness:84 - Average student length: 540.00 +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 20.00 +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_correctness:86 - Length ratio: 27.00 +2025-04-03 at 20:00:10 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-03 at 20:00:10 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.800 Âą 0.173 +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 5.50 Âą 2.81 +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [2, 8, 3, 10, 6, 4] +2025-04-03 at 20:00:10 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The service module first appeared in the docking window at a distance of about 80 feet. The entire bay 4 outer panel was missing, and torn Mylar insulation was seen protruding from the bay. Because of the brilliant reflections from the Mylar, it was difficult to see or photograph any details inside the bay. Initial photography of the service module was conducted through the docking window using the command module 70-mm camera and an 80-mm lens. This camera, the l6-mm sequence camera with a 75-mm lens, and the command module electric still camera with a 250-mm lens were then operated while viewing through the right-hand window. Camera settings were made according to ground instructions. No magazine designation was made by the ground for the sequence camera, so the surface color film was used. +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The transducer is a variable reluctance instrument actuated by differential pressure applied across a twisted Bourdon tube. The housing, including the cavity containing the Bourdon tube and the variable reluctance elements, is nickel plated. The Bourdon tube-variable reluctance + +assembly and the sense port fitting are soldered in place. Inspection of the failed units indicates that the flaking occurs adjacent to the solder. + +The most probable cause of the problem is poor plating adhesion to the aluminum base metal. Differential expansion between the solder and the aluminum may cause the plating to crack. Moisture from the environmental control system suit loop could then penetrate the plating, corrode the aluminum base metal and cause the plating to peel and flake. The nickel flakes could then enter the air gap of the variable reluctance elements and affect the measurement. +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +Command module.- Command module reaction control system propellant usages cannot be accurately assessed, since telemetry data were not available during entry. Until the time of communications blackout, approximately l2 pounds of propellant had been used. For a normal entry, this value would be considered high; however, the system was activated longer than normal and was used during separation from the lunar module. + +Loadedīŧˆ quantities, lb System 1 System 2 Fuel 44.2 44.6 Oxi di zer 77.8 78.5 Totals 122.0 123.1 + +7.1.3 Cryogenic Flui ds + +Cryogenic oxygen and hydrogen usages were nominal until the time of the incident.. The pressure decay in oxygen tank 2 was essentially instantaneous, while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel cells . +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +a. The mission was aborted because of the total loss of primary oxygen in the service module. This loss resulted from an incompatibility between switch design and preflight procedures, a condition which, when combined with an abnormal preflight detanking procedure, caused an inflight shorting and a rapid oxidation within one of two redundant storage tanks. The oxidation then resulted in a loss of pressure integrity in the related tank and eventually in the remaining tank. + +b. The concept of a backup crew was proven for the first time when 3 days prior to flight the backup Command Module Pilot was substituted for his prime-crew counterpart, who was exposed and susceptible to rubella (German measles). + +c. The performance of lunar module systems demonstrated an emergency operational capability. Lunar module systems supported the crew for a period approximately twice their intended design lifetime. +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ +Result 2: +During the peak engine activity period after the oxygen tank incident, engine package temperatures reached as high as 203° F, which is normal for the commanded duty cycles. All reaction control data were normal for the configuration and duty cycles that existed, including the quad C data which showed the system in a nonuse configuration because the isolation valves were closed. System data were normal when checked prior to entry at about l23 hours, at which time the total propellant consumed was 286 pounds (86 pcumds from quad A, 65 from B, 33 from C, and 102 from D). + +5.7.2 Command Module +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ +Result 2: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 6.0 +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 1.000 +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 20:00:10 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.200, Perfect scores: 0/6 +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 2.00 +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 20:00:10 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:00:10 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.564, Max reward: 0.650 +2025-04-03 at 20:00:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:00:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "resolver for shaft axis" +2025-04-03 at 20:00:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 20:00:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Hall effect resolver +2025-04-03 at 20:00:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 20:00:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: potentiometer resolver vs incremental encoder for shaft axis +2025-04-03 at 20:00:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-03 at 20:00:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: resolver axis shaft positioning +2025-04-03 at 20:00:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 20:00:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:00:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vacuum effects on half speed resolver slip rings +2025-04-03 at 20:00:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 20:00:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: angular resolution zero reference +2025-04-03 at 20:00:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ +Result 2: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ + +2025-04-03 at 20:00:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:00:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how does a magnetic resolver with slip ring work +2025-04-03 at 20:00:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 20:00:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: half-speed resolver definition +2025-04-03 at 20:00:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ + +2025-04-03 at 20:00:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:00:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: brushes on half speed resolver +2025-04-03 at 20:00:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 20:00:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: slip ring resolver shaft axis +2025-04-03 at 20:00:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 20:00:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:00:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: half speed resolver vacuum +2025-04-03 at 20:00:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 20:00:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:00:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vacuum effect on half speed resolver +2025-04-03 at 20:00:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 20:00:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:00:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: slip ring half speed resolver +2025-04-03 at 20:00:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ + +2025-04-03 at 20:00:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:00:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: unambiguous zero reference +2025-04-03 at 20:00:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ +Result 2: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ + +2025-04-03 at 20:00:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:00:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: definition of half speed resolver +2025-04-03 at 20:00:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 20:00:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:00:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: half speed resolver application +2025-04-03 at 20:00:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 20:00:28 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:00:28 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:00:28 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, True, False, False] +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_correctness:82 - Student lengths: [170, 112, 232, 639, 756, 1475] +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [19, 19, 19, 19, 19, 19] +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_correctness:84 - Average student length: 564.00 +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 19.00 +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_correctness:86 - Length ratio: 29.68 +2025-04-03 at 20:00:28 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_format:228 - Responses ending properly: 2/6 +2025-04-03 at 20:00:28 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.408 Âą 0.362 +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.67 Âą 3.54 +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [1, 0, 4, 1, 0, 10] +2025-04-03 at 20:00:28 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +A simplified schematic of the optics shaft servo loop mechanization is shown in figure l4-2. In the zero optics mode, the sine outputs of the half-speed and l6-speed resolvers are routed through a coarse/fine switching network to the motor drive amplifier and are used to null the system. Rate feedback from the motor tachometer is routed to the drive amplifier through a compensation network which removes any bias in the signal. When the zero optics mode is selected, the coupling-data-unit counter and the computer register which contains the shaft angle are zeroed for 15 seconds and then released to follow the l6-speed resolver. The half-speed resolver, the fine/coarse switching network, and the tachometer feedback compensation are used only in the zero optics mode. + + + +Figure l4-2.- Zero optics mode circuitry. +------ + +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 20:00:28 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ + +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 20:00:28 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe testing of the half-speed resolver with resistance in the low side of the...', 'Result 1:\nSome evidence of susceptibility to vacuum was exhibited in this class of reso...', 'Result 1:\nSome evidence of susceptibility to vacuum was exhibited in this class of reso...', 'Result 1:\nSome evidence of susceptibility to vacuum was exhibited in this class of reso...'] +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 20:00:28 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ +Result 2: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ + +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ + +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ +Result 2: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ + +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ +Result 2: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ + +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation. + +Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +Some evidence of susceptibility to vacuum was exhibited in this class of resolvers during qualification testing when variations of approximately 5 ohms were observed in the slip ring resistance during thermal + +14-7 + + + +Figure l4-3.- Details of half speed resolver. + + + +Figure l4-4.- One-half speed resolver. + +vacuum testing. The tests were run with the units rotating at l rpm, however, and the momentary resistance changes disappeared with the wiping acti on. +------ +Result 2: +The testing of the half-speed resolver with resistance in the low side of the sine winding and the vacuum susceptibility exhibited during qualification testing closely duplicate the characteristics of inflight "zero optics" operation. The slip-ring mechanism is unique to the shaftaxis, since none of the other resolvers in the system use slip rings. This resolver is in the optics head, which is vented to a vacuum. The rotation of the optics head in a normal operation would wipe the slip rings clean and explain the delay in the fluctuations for some hours after selecting zero optics. + +Corrective action to high resistance on the brush/slip rings of the resolver is not required since accurate zeroing is unaffected and there is no effect in the operation of the system other than system readout when not in use. This condition can be expected to recur in future Apollo flight. Future crews will be briefed on this situation. + +This anomaly is closed. +------ + +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 20:00:28 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 20:00:28 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 20:00:28 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:00:29 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.279, Max reward: 0.680 +2025-04-03 at 20:00:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:00:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens to contaminants when they freeze? +2025-04-03 at 20:00:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 20:00:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens to contaminants when they freeze? +2025-04-03 at 20:00:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 20:00:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens to contaminants when they freeze +2025-04-03 at 20:00:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 20:00:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what happens to contaminant particles when they freeze? +2025-04-03 at 20:00:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 20:00:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens to contaminants when they freeze? +2025-04-03 at 20:00:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 20:00:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:00:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens to contaminants when they freeze in a closed system? +2025-04-03 at 20:00:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 20:00:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: chemical changes in contaminants when they freeze +2025-04-03 at 20:00:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 20:00:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens to gas bubbles and pressure when different types of contaminants freeze in a spacecraft or potable water system? +2025-04-03 at 20:00:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 20:00:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:00:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: how contaminants change when they freeze +2025-04-03 at 20:00:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 20:00:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the effect of freezing on nickel and bacterial contaminants in space-related water systems? +2025-04-03 at 20:00:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 20:00:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:00:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effects of freezing on gases on surfaces +2025-04-03 at 20:00:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 20:00:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens to nickel, a contaminant that is dissolved in water, when it freezes in pipes? +2025-04-03 at 20:00:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 20:00:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:00:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effects of freezing on groundwater contaminants +2025-04-03 at 20:00:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 20:00:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens to nickel dissolved in water when it freezes in pipe systems? +2025-04-03 at 20:00:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 20:00:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:00:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what happens to contaminants frozen in ground water +2025-04-03 at 20:00:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 20:00:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Does freezing affect the dissolved nickel in water in pipe systems? +2025-04-03 at 20:00:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 20:00:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:00:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effects of freezing on suspended solids in water +2025-04-03 at 20:00:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:00:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Can dissolved nickel remain stable in water even after freezing and gas bubbles formation in pipes? +2025-04-03 at 20:00:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 20:00:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:00:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Does nickel in dissolved form remain stable in frozen, pressurized pipes? +2025-04-03 at 20:00:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 20:00:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:00:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens to dissolved nickel in water when it freezes and condenses on pipe surfaces? +2025-04-03 at 20:00:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 20:00:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:00:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What happens to dissolved nickel when it condenses on pipe surfaces in a frozen state? +2025-04-03 at 20:00:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 20:00:51 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:00:51 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:00:51 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, True, False, False] +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_correctness:82 - Student lengths: [567, 386, 480, 868, 896, 250] +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [92, 92, 92, 92, 92, 92] +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_correctness:84 - Average student length: 574.50 +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 92.00 +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_correctness:86 - Length ratio: 6.24 +2025-04-03 at 20:00:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 20:00:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.517 Âą 0.341 +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 3.50 Âą 3.69 +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [2, 7, 1, 1, 10, 0] +2025-04-03 at 20:00:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7. + +All available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation. + +5.8 ENVIRONMENIAL CONTROL +------ + +2025-04-03 at 20:00:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nfreezing in the lines most likely occurred. This freezing condition could hav...', 'Result 1:\nfreezing in the lines most likely occurred. This freezing condition could hav...'] +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 20:00:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nfreezing in the lines most likely occurred. This freezing condition could hav...'] +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 20:00:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...'] +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ + +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ + +2025-04-03 at 20:00:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nfreezing in the lines most likely occurred. This freezing condition could hav...', 'Result 1:\nfreezing in the lines most likely occurred. This freezing condition could hav...', 'Result 1:\nPreflight testing of both command module and lunar module water supplies reve...', 'Result 1:\nPreflight testing of both command module and lunar module water supplies reve...', 'Result 1:\nfreezing in the lines most likely occurred. This freezing condition could hav...', 'Result 1:\nfreezing in the lines most likely occurred. This freezing condition could hav...', 'Result 1:\nfreezing in the lines most likely occurred. This freezing condition could hav...', 'Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...', 'Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...', 'Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...'] +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +2025-04-03 at 20:00:51 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 1.0 +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.167 +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 20:00:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.133, Perfect scores: 0/6 +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 20:00:51 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:00:51 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.327, Max reward: 0.617 +2025-04-03 at 20:00:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:00:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: saturn ib launch date +2025-04-03 at 20:00:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 20:00:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query: first Lunar Module mission date +2025-04-03 at 20:00:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 20:00:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 4 uncrewed Saturn IV B launch date +2025-04-03 at 20:00:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:00:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 4 mission launch date +2025-04-03 at 20:00:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 20:00:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 launch date +2025-04-03 at 20:00:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 20:00:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:00:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 launch date april 11, 1970 +2025-04-03 at 20:00:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 20:00:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 4 spacecraft Saturn IV B impact point +2025-04-03 at 20:00:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 20:00:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 4 launch date April 1968 +2025-04-03 at 20:00:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:00:58 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 launch march 1967 +2025-04-03 at 20:00:58 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:00:58 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:01:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 launch date kennedy space center april 11 1970 +2025-04-03 at 20:01:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 20:01:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 4 S-IVB impact signal explanation +2025-04-03 at 20:01:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:01:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 4 launch +2025-04-03 at 20:01:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 20:01:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: starfire AA-4 launch date +2025-04-03 at 20:01:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 20:01:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:01:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 4 S-IVB impact explanation initial rapid buildup +2025-04-03 at 20:01:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 20:01:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 launch date november 1967 +2025-04-03 at 20:01:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:01:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:01:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: unexplained S-IVB impact signal components +2025-04-03 at 20:01:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 20:01:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: projects using satellite assist Apollo 4 +2025-04-03 at 20:01:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 20:01:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:01:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 4 mission launch date +2025-04-03 at 20:01:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 20:01:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: april 1967 apollo 4 +2025-04-03 at 20:01:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:01:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:01:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 4 launch date +2025-04-03 at 20:01:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 20:01:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 4 launch site Kennedy Space Center +2025-04-03 at 20:01:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 20:01:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:01:10 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:01:10 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:01:11 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, True, False, True] +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_correctness:82 - Student lengths: [513, 343, 60, 403, 141, 209] +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [12, 12, 12, 12, 12, 12] +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_correctness:84 - Average student length: 278.17 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 12.00 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_correctness:86 - Length ratio: 23.18 +2025-04-03 at 20:01:11 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 20:01:11 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.592 Âą 0.335 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 3.50 Âą 2.69 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [3, 1, 7, 0, 3, 7] +2025-04-03 at 20:01:11 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 20:01:11 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...'] +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-03 at 20:01:11 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 20:01:11 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.167, Perfect scores: 1/6 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.00 +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 20:01:11 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:01:11 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.435, Max reward: 0.692 +2025-04-03 at 20:01:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:01:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "history of weather photography" or " duration of earth observation" +2025-04-03 at 20:01:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 20:01:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Earth climate photography history +2025-04-03 at 20:01:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ + +2025-04-03 at 20:01:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Earth climate history duration photography +2025-04-03 at 20:01:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ + +2025-04-03 at 20:01:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "length of earth's weather patterns" or "duration of earth's climate" +2025-04-03 at 20:01:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 20:01:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How long were the NASA's Terra and Aqua missions active +2025-04-03 at 20:01:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 20:01:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:01:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: early Apollo photography mission +2025-04-03 at 20:01:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 20:01:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: duration of Earth weather photography project +2025-04-03 at 20:01:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 20:01:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "diffeable atmospheric circulation" or "earth weather duration historical analysis" +2025-04-03 at 20:01:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ +Result 2: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ + +2025-04-03 at 20:01:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How long was the NASA's Earth Observing Systems mission active +2025-04-03 at 20:01:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 20:01:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:01:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ground testing photography during Apollo missions +2025-04-03 at 20:01:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-03 at 20:01:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: when did translunar injection start +2025-04-03 at 20:01:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 20:01:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "earth weather patterns duration spacecraft orbit" +2025-04-03 at 20:01:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ + +2025-04-03 at 20:01:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Is there a specific time frame for the "earth weather photography" task in the Apollo 13 mission +2025-04-03 at 20:01:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-03 at 20:01:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:01:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 10 cloud height photography experiment +2025-04-03 at 20:01:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 20:01:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo translunar injection time +2025-04-03 at 20:01:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 20:01:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo mission earth weather photography duration" or "long duration earth weather patterns" +2025-04-03 at 20:01:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-03 at 20:01:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How long did the earth weather photography test last +2025-04-03 at 20:01:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 20:01:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:01:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar photography duration +2025-04-03 at 20:01:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +urine trans. fer line. Also included was a lunar topographic camera, which could be installed in the command module hatch window for high resolution photography of the lunar surface from orbit. The camera provided a 4.5-inch film format and had an 18-inch focal length and image-motion compensation. The photographs would yield a resolution of approximately l2 feet and would include a l5-mile square area on the surface for each frame exposed. +------ + +2025-04-03 at 20:01:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 translunar injection time +2025-04-03 at 20:01:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 20:01:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo 6 earth weather photography duration" or " duration of apollo 6 mission earth weather pattern observation" +2025-04-03 at 20:01:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 20:01:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the duration of the incompatible information and can it be validated +2025-04-03 at 20:01:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 20:01:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:01:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar surface photography characteristics +2025-04-03 at 20:01:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +urine trans. fer line. Also included was a lunar topographic camera, which could be installed in the command module hatch window for high resolution photography of the lunar surface from orbit. The camera provided a 4.5-inch film format and had an 18-inch focal length and image-motion compensation. The photographs would yield a resolution of approximately l2 feet and would include a l5-mile square area on the surface for each frame exposed. +------ + +2025-04-03 at 20:01:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Aborted orbital translunar injection +2025-04-03 at 20:01:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:01:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo 6 mission duration" +2025-04-03 at 20:01:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 20:01:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the duration of earth weather photography +2025-04-03 at 20:01:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-03 at 20:01:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:01:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar surface photography duration +2025-04-03 at 20:01:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +urine trans. fer line. Also included was a lunar topographic camera, which could be installed in the command module hatch window for high resolution photography of the lunar surface from orbit. The camera provided a 4.5-inch film format and had an 18-inch focal length and image-motion compensation. The photographs would yield a resolution of approximately l2 feet and would include a l5-mile square area on the surface for each frame exposed. +------ + +2025-04-03 at 20:01:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "apollo 6 mission duration earth weather photography" +2025-04-03 at 20:01:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-03 at 20:01:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:01:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: duration of photography during Apollo missions +2025-04-03 at 20:01:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 20:01:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "duree de l'atmosphere" +2025-04-03 at 20:01:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 ArcturugīŧŒ36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ + +2025-04-03 at 20:01:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:01:34 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:01:34 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:01:34 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, True, False, False, True, False] +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_correctness:82 - Student lengths: [414, 814, 185, 1758, 238, 325] +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [21, 21, 21, 21, 21, 21] +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_correctness:84 - Average student length: 622.33 +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 21.00 +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_correctness:86 - Length ratio: 29.63 +2025-04-03 at 20:01:34 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 20:01:34 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.692 Âą 0.379 +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 4.83 Âą 3.18 +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [1, 8, 6, 8, 6, 0] +2025-04-03 at 20:01:34 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo īŧ‰11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +urine trans. fer line. Also included was a lunar topographic camera, which could be installed in the command module hatch window for high resolution photography of the lunar surface from orbit. The camera provided a 4.5-inch film format and had an 18-inch focal length and image-motion compensation. The photographs would yield a resolution of approximately l2 feet and would include a l5-mile square area on the surface for each frame exposed. +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +urine trans. fer line. Also included was a lunar topographic camera, which could be installed in the command module hatch window for high resolution photography of the lunar surface from orbit. The camera provided a 4.5-inch film format and had an 18-inch focal length and image-motion compensation. The photographs would yield a resolution of approximately l2 feet and would include a l5-mile square area on the surface for each frame exposed. +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ +Result 2: +urine trans. fer line. Also included was a lunar topographic camera, which could be installed in the command module hatch window for high resolution photography of the lunar surface from orbit. The camera provided a 4.5-inch film format and had an 18-inch focal length and image-motion compensation. The photographs would yield a resolution of approximately l2 feet and would include a l5-mile square area on the surface for each frame exposed. +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 20:01:34 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nTo aid in a test of the feasibility of performing stereoscopic determination ...', 'Result 1:\nSupplement number Title Publication date/status Apollo 10 1 Trajectory Recons...', 'Result 1:\nDescription Completed B C Television coverage No Contingency sample collectio...', 'Result 1:\nTo aid in a test of the feasibility of performing stereoscopic determination ...', 'Result 1:\nDescription Completed B C Television coverage No Contingency sample collectio...', 'Result 1:\nDescription Completed B C Television coverage No Contingency sample collectio...', 'Result 1:\nDescription Completed B C Television coverage No Contingency sample collectio...', 'Result 1:\nTo aid in a test of the feasibility of performing stereoscopic determination ...'] +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops. + +This anomaly is closed. + +15.0 CONCLUS IONS + +The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ +Result 2: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +MSC-02680 + +CHANGE SHEET + +FOR + +NASA-MSC INTERNAL REPORT + +APOLLO 13 MISSION REPORT + +Change 1 + + + +May 1970 + +James A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program + +After the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover "Change l inserted." + +In addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated. + +NOTE: A black bar in the margin of affected pages indicates the information that was changed or added. + +7.1.6 Batteries +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The determination of the wind field in the atmosphere is one of the prime requirements for accurate long-range numerical weather prediction. Wind fields are also the most difficult to measure with the desired sample density (as discussed in ref. 4)." The output of the geosynchronous Advanced Technology Satellites I and IIl is now being used as a crude estimate of wind fields by comparing the translation of clouds between successive frames 20 minutes apart. This comparison does not define the wind field, however, as a function of height above the surface, which is an important restriction to data application. The ability to determine the height of cloud elements would add this dimension to the satellite wind field analysis. A capability to determine cloud height has been demonstrated by use of stereographic photogrammetry on low altitude photographs taken from Apollo 6 (ref. 5). This success suggests that cloud heights and therefore wind velocity may also be determined by using data +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY + +Magazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130°00'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28°25'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27°39'N 151°39*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156°35'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27°14'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27°04+N 165°9*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26°54'N 170°50'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175°51'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26°36'N 179°14*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26°27'N g60 56728 9.251 1.6254 2436.6 +------ +Result 2: +Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 ArcturugīŧŒ36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263 + +preferred alignment DRererence metrix (REFSMAT) CCoarse alignment +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST ¡¡¡¡¡ 8-11 8.10 ENTRY AND LANDING.¡¡. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY ¡¡¡¡¡¡¡ ¡ 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL ¡¡¡¡ 10-1 10.2 NETWORKīŧŽÂˇÂˇÂˇÂˇÂˇÂˇÂˇ 10-2 10.3 RECOVERY OPERATIONS...¡.¡. ¡¡¡ 10-2 11.0 EXPERIMENTS¡¡¡¡¡īŧŽÂˇÂˇÂˇ.¡¡¡¡. ¡¡ 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITESīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇīŧŽÂˇÂˇīŧŽÂˇÂˇ 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . ¡¡ 12-1 13.0 LAUNCH VEHICLE SUMMARY¡......¡......... 13-1 14.0 ANOMALY SUMMARY ¡¡¡¡¡¡¡¡¡¡ 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . ¡¡ 14-1 14.2 LUNAR MODULE ¡¡¡¡¡¡¡¡¡¡¡¡ 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT ¡¡¡ 14-36 15.0 CONCLUSIONS¡¡¡¡.¡¡¡.¡¡īŧŽÂˇ.īŧŽÂˇÂˇ ¡¡ 15-1 APPENDIX A - VEHICLE +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +approximately l mile of the target location. Because of power restrictions, the circuit breaker for the data storage equipment recorder was left open during entry, and no entry data are available for an entry performance analysis . +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ +Result 2: +To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately, +------ + +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +2025-04-03 at 20:01:34 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 20:01:34 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.167, Perfect scores: 0/6 +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.83 +2025-04-03 at 20:01:34 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 20:01:34 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:01:35 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.281, Max reward: 0.700 +2025-04-03 at 20:01:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:01:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Manned Spacecraft Center location +2025-04-03 at 20:01:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "NASA Manned Spacecraft Center location +2025-04-03 at 20:01:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Manned Spacecraft Center location +2025-04-03 at 20:01:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa johannesburg space center location +2025-04-03 at 20:01:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 20:01:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa manned spacecraft center location +2025-04-03 at 20:01:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:01:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reliable sources for NASA Manned Spacecraft Center location +2025-04-03 at 20:01:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 20:01:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Manned Spacecraft Center contact information +2025-04-03 at 20:01:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location nasa houston texas +2025-04-03 at 20:01:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 20:01:40 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Manned Spacecraft Center Houston, Texas +2025-04-03 at 20:01:40 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:01:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Manned Spacecraft Center current location +2025-04-03 at 20:01:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Manned Spacecraft Center zip code Houston +2025-04-03 at 20:01:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 20:01:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: nasa houston texas address +2025-04-03 at 20:01:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 20:01:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:01:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Houston NASA Manned Spacecraft Center facility location +2025-04-03 at 20:01:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 20:01:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Manned Spacecraft Center exact location 77058 +2025-04-03 at 20:01:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:01:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Johnson Space Center Houston +2025-04-03 at 20:01:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 20:01:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:01:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: NASA Houston Texas center +2025-04-03 at 20:01:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 20:01:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:01:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Johnson Space Center location +2025-04-03 at 20:01:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:01:50 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:01:50 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:01:50 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, True, True, False, False] +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_correctness:82 - Student lengths: [266, 396, 285, 276, 71, 280] +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [14, 14, 14, 14, 14, 14] +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_correctness:84 - Average student length: 262.33 +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 14.00 +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_correctness:86 - Length ratio: 18.74 +2025-04-03 at 20:01:50 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 20:01:50 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.542 Âą 0.310 +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 2.83 Âą 2.27 +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [1, 7, 4, 3, 2, 0] +2025-04-03 at 20:01:50 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:50 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...'] +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ + +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:50 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...', 'Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...', 'Result 1:\nPOSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION\n\nNASA-Mann...', 'Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...'] +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION + +NASA-Manned Spacecraft Center Houston, Texas 77058 + +ATTN: PT2(office Symbol) + +(Continued from inside front cover) +------ + +2025-04-03 at 20:01:50 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...', 'Result 1:\nMSC-02680\n\nDISTRIBUTION AND REFERENCING\n\nThis paper is not suitable for gener...'] +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +2025-04-03 at 20:01:50 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 2.0 +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.333 +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 20:01:50 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.567, Perfect scores: 2/6 +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 5/6 +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.83 +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 3/6 +2025-04-03 at 20:01:50 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:01:50 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.261, Max reward: 0.636 +2025-04-03 at 20:01:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:01:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: struct area oscillations S-II +2025-04-03 at 20:01:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 20:01:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "LHC S-II crossbeam oscillations peak amplitude" +2025-04-03 at 20:01:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:01:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-II crossbeam oscillations peak amplitude 33.7g time +2025-04-03 at 20:01:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:01:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Carl Sagan S-II crossbeam oscillations +33.7g time +2025-04-03 at 20:01:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 20:01:56 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "S-II oscillations crossbeam experiment 1960s peak 33.7g" +2025-04-03 at 20:01:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:01:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:01:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-II crossbeam oscillations peak amplitude +33.7g time +2025-04-03 at 20:01:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:01:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("Apollo 13 S-II crossbeam oscillation peak amplitude" OR "Apollo 13 S-II Gimbals expected maximum amplitude") +2025-04-03 at 20:01:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 20:01:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-II crossbeam oscillations peak amplitude 33.7g +2025-04-03 at 20:01:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 20:01:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 S-II crossbeam oscillations +33.7g +2025-04-03 at 20:01:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-03 at 20:01:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 13 S-II crossbeam oscillations +33.7g peak time" +2025-04-03 at 20:01:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-03 at 20:01:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:02:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 S-II crossbeam oscillations time +2025-04-03 at 20:02:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-03 at 20:02:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("Apollo 13 S-II oscillations peak time" OR "Apollo 13 emergency acceleration timing") +2025-04-03 at 20:02:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 20:02:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 mission S-IVB impact time +2025-04-03 at 20:02:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 20:02:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 S-II stage 5 early shutdown +33.7g +2025-04-03 at 20:02:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 20:02:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: none needed, as the answer can be directly inferred from the provided information +2025-04-03 at 20:02:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +sufficient accuracy to permit a safe earth entry. +------ +Result 2: +according to ground instructions. +------ + +2025-04-03 at 20:02:02 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:02:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("Apollo 13 S-II oscillations peak time" OR "Apollo 13 S-II engine vibration timing") +2025-04-03 at 20:02:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 20:02:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-II crossbeam oscillations timing after S-IVB impact +2025-04-03 at 20:02:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:02:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 engine 5 shutdown oxygen tank 2 loss of power +2025-04-03 at 20:02:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 20:02:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: none needed, no specific information about orbital parameters +2025-04-03 at 20:02:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-03 at 20:02:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:02:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("Apollo 13 S-II oscillation peak time" OR "Apollo 13 S-II vibration timing") +2025-04-03 at 20:02:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 20:02:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 timing of S-II oscillations before S-IVB impact during lift-off +2025-04-03 at 20:02:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 20:02:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 crew warning system cryogenic tank incident +2025-04-03 at 20:02:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:02:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: none needed, as the answer can be directly inferred from the provided information +2025-04-03 at 20:02:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +sufficient accuracy to permit a safe earth entry. +------ +Result 2: +according to ground instructions. +------ + +2025-04-03 at 20:02:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:02:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("Apollo 13 S-II oscillations timing relative to cue time" OR "Apollo 13 engine vibration timing after S-II ignition") +2025-04-03 at 20:02:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 20:02:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-II oscillations during lift-off amplitude and frequency +2025-04-03 at 20:02:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-03 at 20:02:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module cold soak venting +2025-04-03 at 20:02:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 20:02:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 nominal re-entry conditions +2025-04-03 at 20:02:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 20:02:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:02:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ("Apollo 13 S-II boost timing of max amplitude 33.7g" OR "Apollo 13 S-II vibration timing 33.7g") +2025-04-03 at 20:02:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:02:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-II crossbeam oscillations time during S-IC boost +2025-04-03 at 20:02:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:02:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 gas leak docking timnel insulation +2025-04-03 at 20:02:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:02:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission trajectory +2025-04-03 at 20:02:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 20:02:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:02:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-II crossbeam oscillations to S-IVB impact +2025-04-03 at 20:02:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 20:02:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 helium pressure lunar module descent propulsion +2025-04-03 at 20:02:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 20:02:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 geology field trip +2025-04-03 at 20:02:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 20:02:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:02:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 time difference between S-II oscillations and S-IVB impact +2025-04-03 at 20:02:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 20:02:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module powered-down electrical power +2025-04-03 at 20:02:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 20:02:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 seismic signals +2025-04-03 at 20:02:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 20:02:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:02:19 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 S-IVB impact signals explanation +2025-04-03 at 20:02:19 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:02:20 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:02:20 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:02:20 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, True, False, False, False, False] +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_correctness:82 - Student lengths: [103, 256, 1862, 341, 1983, 1809] +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [13, 13, 13, 13, 13, 13] +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_correctness:84 - Average student length: 1059.00 +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 13.00 +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_correctness:86 - Length ratio: 81.46 +2025-04-03 at 20:02:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-03 at 20:02:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.621 Âą 0.361 +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 7.17 Âą 4.37 +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 2/6 +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [4, 7, 9, 0, 9, 14] +2025-04-03 at 20:02:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 20:02:20 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of Âą225 psi initiated engine cutoff through the "thrust OK" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +sufficient accuracy to permit a safe earth entry. +------ +Result 2: +according to ground instructions. +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +sufficient accuracy to permit a safe earth entry. +------ +Result 2: +according to ground instructions. +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation + +REFERENCES + +Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970. + +Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970. +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 5.0 +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.833 +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 20:02:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.467, Perfect scores: 1/6 +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 4/6 +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 2.50 +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 20:02:20 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:02:20 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.351, Max reward: 0.783 +2025-04-03 at 20:02:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:02:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: extraneous firings in experiments +2025-04-03 at 20:02:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-03 at 20:02:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the effects of repeated firings of a machine on a particular device? +2025-04-03 at 20:02:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-03 at 20:02:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: extraneous firings figure 5.6-1 adverse effects +2025-04-03 at 20:02:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 20:02:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Extraneous electrical firings effects v.s electrical arcing effects +2025-04-03 at 20:02:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ + +2025-04-03 at 20:02:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "plastic deformation instability thermal stress" +2025-04-03 at 20:02:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At approximately lo9 hours, the burst disk in the supercritical helium tank ruptured, as expected. The venting caused an unexpected reversal in the lunar module yaw rate [command module roll] during passive thermal control at about twice the initial value and also introduced some pitch motion. No attempt was made, however, to reestablish manually a stable passive thermal control mode. + +8.9.2 Entry Preparation +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-03 at 20:02:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:02:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission 5-70-5825 extraneous firings +2025-04-03 at 20:02:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:02:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What were the effects of extraneous firings on the thermal control system in the Apollo Guidance Computer or a similar system? +2025-04-03 at 20:02:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:02:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacecraft extraneous firings figure 5.6-1 thermal control +2025-04-03 at 20:02:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 20:02:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn rocket B083 launch electrical arcing effects +2025-04-03 at 20:02:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 20:02:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "lunar module helicopters temperature sensors" +2025-04-03 at 20:02:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 20:02:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:02:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 extraneous firings SIC and S-II +2025-04-03 at 20:02:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 20:02:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What effects do extraneous firings have on spacecraft performance or systems? +2025-04-03 at 20:02:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 20:02:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 launch electrical disturbances effects Saturn rocket +2025-04-03 at 20:02:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 20:02:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "thermocouple failure temperatures" +2025-04-03 at 20:02:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 20:02:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:02:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 engine issues SIC S-II +2025-04-03 at 20:02:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:02:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the effects of extraneous firings on pyrotechnic systems in spacecraft? +2025-04-03 at 20:02:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-03 at 20:02:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: effects of launch voltage separation on Apollo 12 and 13 Saturn rocket +2025-04-03 at 20:02:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-03 at 20:02:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "command module ice accumulation and heat transfer" +2025-04-03 at 20:02:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ + +2025-04-03 at 20:02:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:02:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module descent propulsion system pressure concerns +2025-04-03 at 20:02:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:02:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the effects of extraneous firings on spacecraft systems, specifically thermal control modes? +2025-04-03 at 20:02:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-03 at 20:02:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 and 13 Saturn rocket electrical equilibrium potentials +2025-04-03 at 20:02:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:02:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "diesel to cryogenic pressure and fuel difference" +2025-04-03 at 20:02:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-03 at 20:02:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:02:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 helium cooling system pressurization issues +2025-04-03 at 20:02:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:02:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the long-term effects of extraneous firings on spacecraft thermal control modes? +2025-04-03 at 20:02:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-03 at 20:02:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V electrostatic potential measurements +2025-04-03 at 20:02:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:02:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "cryogenic tank impact on fuel cell pressure" +2025-04-03 at 20:02:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ +Result 2: +Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal. + +During the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l. +------ + +2025-04-03 at 20:02:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:02:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium cooling system functionality engine performance Apollo 13 +2025-04-03 at 20:02:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:02:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the immediate effects of extraneous firings on spacecraft systems, specifically those related to attitude control? +2025-04-03 at 20:02:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ + +2025-04-03 at 20:02:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V electrostatic potential and ground measurements +2025-04-03 at 20:02:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:02:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "operating effects of un committed diesel fuel pressure loss" +2025-04-03 at 20:02:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 20:02:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:02:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the response of the digital autopilot to unauthorized firings or system failures in spacecraft? +2025-04-03 at 20:02:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-03 at 20:02:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V electrostatic potential and ground measurements +2025-04-03 at 20:02:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:02:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:03:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the consequences of automatic systems failing to re-initialize after a critical system failure in spacecraft? +2025-04-03 at 20:03:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ + +2025-04-03 at 20:03:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V electrostatic potential & ground measurements +2025-04-03 at 20:03:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:03:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:03:02 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the overall impact of minor errors or glitches in a networked spacecraft system on its overall performance and control? +2025-04-03 at 20:03:02 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-03 at 20:03:02 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:03:02 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:03:02 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, False, True, True, False] +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_correctness:82 - Student lengths: [509, 776, 1512, 1071, 185, 1655] +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [41, 41, 41, 41, 41, 41] +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_correctness:84 - Average student length: 951.33 +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 41.00 +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_correctness:86 - Length ratio: 23.20 +2025-04-03 at 20:03:02 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 20:03:02 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.558 Âą 0.326 +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 22.83 Âą 38.27 +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 2/6 +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 7, 10, 2, 10, 108] +2025-04-03 at 20:03:02 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-03 at 20:03:02 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +------ +Result 2: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +------ +Result 2: +To establish a passive thermal control mode during initial transearth coast, the spacecraft was manually maneuvered to the initial attitude by nulling out the attitude error needles. In this position, spacecraft. rates were monitored by the ground. When rates were sufficiently damped, 2l yaw-right pulse inputs were made to establish a vehicle rolling motion. The resulting maneuver placed the apparent moon and earth motion horizontal with respect to the lunar module windows. + +After the passive thermal control mode was established, the lumar module was powered down according to the contingency checklist for an emergency power-down. Minor modifications were made to this procedure to account for passive thermal control mode operation. The spacecraft functions remaining were low-bit-rate telemetry, S-band tracking and voice, caution and warming sensing, cabin repressurization capability, and the operation of the glycol pumps and suit fans. +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. + +The above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood. +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:03:02 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe next series of events occurred within a fraction of a second between the ...', 'Result 1:\nAs a result of the electrical disturbances experienced during the Apollo l2 l...', 'Result 1:\nAs a result of the electrical disturbances experienced during the Apollo l2 l...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...'] +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally. + + + +NASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode. +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +At approximately lo9 hours, the burst disk in the supercritical helium tank ruptured, as expected. The venting caused an unexpected reversal in the lunar module yaw rate [command module roll] during passive thermal control at about twice the initial value and also introduced some pitch motion. No attempt was made, however, to reestablish manually a stable passive thermal control mode. + +8.9.2 Entry Preparation +------ +Result 2: +The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform + +alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed. + +At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground. + +8.7.2 First Midcourse Correction +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ +Result 2: +Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal. + +During the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l. +------ + +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 20:03:02 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nAt approximately lo9 hours, the burst disk in the supercritical helium tank r...', 'Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...', 'Result 1:\nAll mechanical systems functioned properly. One mechanical anomaly, however, ...', 'Result 1:\nDuring the period when the command module was powered down, the cabin tempera...', 'Result 1:\nThe supercritical helium pressurization system displayed abnormal performance...', 'Result 1:\nUnusual variations in the oxygen flow rates to all three fuel cells were obse...', 'Result 1:\nFollowing recovery of the data, the vehicle had experienced a translation cha...'] +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 20:03:02 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.100, Perfect scores: 0/6 +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.17 +2025-04-03 at 20:03:02 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 20:03:02 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:03:03 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.353, Max reward: 0.855 +2025-04-03 at 20:03:04 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:03:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB engine cutoff time +2025-04-03 at 20:03:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 20:03:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-ivb engine cutoff time +2025-04-03 at 20:03:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 20:03:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB engine cutoff event +2025-04-03 at 20:03:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 20:03:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo SIVB engine cutoff +2025-04-03 at 20:03:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-03 at 20:03:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB engine cutoff Mission 7 Apollo +2025-04-03 at 20:03:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:03:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB engine cutoff date +2025-04-03 at 20:03:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 20:03:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:03:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB engine cutoff time Apollo +2025-04-03 at 20:03:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:03:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: compare s-ivb engine cutoff times +2025-04-03 at 20:03:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 20:03:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 S-IVB engine cutoff exact time +2025-04-03 at 20:03:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 20:03:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ap10ll SIVB engine cutoff time +2025-04-03 at 20:03:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 20:03:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 S-IVB engine cutoff time +2025-04-03 at 20:03:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 20:03:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:03:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB engine cutoff time lunar impact +2025-04-03 at 20:03:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 20:03:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-ivb s-ii engine cutoff time apollo 12 +2025-04-03 at 20:03:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:03:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Ap10ll sivb inner engine cutfty +2025-04-03 at 20:03:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 20:03:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 00:12:30 S-IVB engine cutoff +2025-04-03 at 20:03:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 20:03:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:03:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB impact time successful lunar landing +2025-04-03 at 20:03:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 20:03:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-ivb auxiliary propulsion system firing time s-ivb engine cutoff +2025-04-03 at 20:03:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 20:03:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the correct SIVB engine cutoff time +2025-04-03 at 20:03:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 20:03:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +2025-04-03 at 20:03:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 20:03:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:03:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-ivb ignition times midcourse correction +2025-04-03 at 20:03:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 20:03:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +2025-04-03 at 20:03:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 20:03:19 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:03:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-ivb ignition times compare to 00:12:30 +2025-04-03 at 20:03:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 20:03:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:03:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: discarded s-ivb ignition time hours minutes seconds +2025-04-03 at 20:03:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 20:03:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:03:24 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-ivb ignition times compare 00:12:30 +2025-04-03 at 20:03:24 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 20:03:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:03:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-ivb ignition time hour value 00 compared to 12 +2025-04-03 at 20:03:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 20:03:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:03:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: s-ivb ignition time 2 hours 0 minutes 30 seconds +2025-04-03 at 20:03:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 20:03:27 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:03:27 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:03:27 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, True, True, True] +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.67 +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_correctness:82 - Student lengths: [208, 905, 221, 365, 64, 494] +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [8, 8, 8, 8, 8, 8] +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_correctness:84 - Average student length: 376.17 +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 8.00 +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_correctness:86 - Length ratio: 47.02 +2025-04-03 at 20:03:27 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.667, Valid formats: 4.0/6 +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-03 at 20:03:27 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.721 Âą 0.232 +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 4.67 Âą 2.56 +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [4, 10, 2, 4, 3, 5] +2025-04-03 at 20:03:27 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ +Result 2: +The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +The planned launch and earth parking orbit phases for this mission were very similar to those for Apollo l2. However, during the second stage (S-Il) boost into the planned l00-mile circular parking orbit, the center engine cut off about l32 seconds early ard caused the remaining four engines to burn approximately 34 seconds longer than predicted (as discussed in section l3.0 and reference 2). Space vehicle velocity after S-II boost was 223 ft/sec lower than planned, and as a result, the S-IVB orbital insertion maneuver was approximately 9 seconds longer than predicted, with cutoff velocity within about l.2 ft/sec of the planned value. The total time to orbital insertion was about 44 seconds longer than predicted, with actual. parking orbit parameters of l00.2 by 98.0 miles. +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8 +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +(section ll.3). +------ +Result 2: ++21.0 +3.0 +------ + +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 6.0 +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 1.000 +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 20:03:27 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.567, Perfect scores: 2/6 +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 4/6 +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.67 +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 3/6 +2025-04-03 at 20:03:27 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:03:27 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.588, Max reward: 0.674 +2025-04-03 at 20:03:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:03:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 mission publications +2025-04-03 at 20:03:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 20:03:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 mission reports scientific journal articles +2025-04-03 at 20:03:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 20:03:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission reports by NASA +2025-04-03 at 20:03:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 20:03:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:03:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 mission reports +2025-04-03 at 20:03:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:03:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 supplemental reports +2025-04-03 at 20:03:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 20:03:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission document count and types +2025-04-03 at 20:03:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 20:03:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:03:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 mission report documents +2025-04-03 at 20:03:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 20:03:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 table e-i +2025-04-03 at 20:03:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 20:03:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission documents filtered by type and date +2025-04-03 at 20:03:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 20:03:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:03:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 mission report +2025-04-03 at 20:03:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 20:03:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 supplemental reports table +2025-04-03 at 20:03:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:03:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission total document count +2025-04-03 at 20:03:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 20:03:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:03:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 report mission summary +2025-04-03 at 20:03:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 20:03:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: total number of apollo 8 mission reports +2025-04-03 at 20:03:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 20:03:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 mission evaluations and analysis +2025-04-03 at 20:03:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 20:03:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:03:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 8 technical report +2025-04-03 at 20:03:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 20:03:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 crew training document +2025-04-03 at 20:03:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:03:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:03:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: official apollo 8 mission summary +2025-04-03 at 20:03:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 20:03:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 crew training evaluation report +2025-04-03 at 20:03:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:03:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:03:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Text of official Apollo 8 mission report +2025-04-03 at 20:03:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 20:03:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 training evaluation procedures +2025-04-03 at 20:03:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 20:03:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:03:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: official nasa apollo 8 mission report +2025-04-03 at 20:03:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 20:03:54 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:03:54 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:03:55 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_correctness:82 - Student lengths: [2033, 472, 65, 366, 2058, 438] +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [1, 1, 1, 1, 1, 1] +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_correctness:84 - Average student length: 905.33 +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 1.00 +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_correctness:86 - Length ratio: 905.33 +2025-04-03 at 20:03:55 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 20:03:55 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.475 Âą 0.475 +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 3.67 Âą 3.86 +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [9, 5, 0, 0, 8, 0] +2025-04-03 at 20:03:55 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-03 at 20:03:55 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-03 at 20:03:55 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +2025-04-03 at 20:03:55 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 20:03:55 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.267, Perfect scores: 1/6 +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.67 +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 20:03:55 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:03:55 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.212, Max reward: 0.577 +2025-04-03 at 20:03:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:04:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: as-201 sc-009 supracircular entry february 26 1966 +2025-04-03 at 20:04:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ + +2025-04-03 at 20:04:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 entry February 26 1966 location +2025-04-03 at 20:04:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 20:04:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 Supercircular entry site February 26 1966 +2025-04-03 at 20:04:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-03 at 20:04:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 Supercircular entry location February 26 1966 +2025-04-03 at 20:04:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-03 at 20:04:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 Supercircular entry February 26 1966 location +2025-04-03 at 20:04:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-03 at 20:04:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:04:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of AS-201 SC-009 Supercircular entry February 26 1966 near N. Mex. +2025-04-03 at 20:04:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-03 at 20:04:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 launch site and entry conditions February 26, 1966 +2025-04-03 at 20:04:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 20:04:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 Supercircular entry February 26 1966 location launch site +2025-04-03 at 20:04:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 20:04:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:04:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 Supercircular entry February 26 1966 exact location north of the US-Mexico border +2025-04-03 at 20:04:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 20:04:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Kennedy Space Center launch site AS-201 SC-009 Supercircular entry +2025-04-03 at 20:04:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 20:04:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: White Sands Missile Range, AS-201 SC-009 Supercircular entry February 26, 1966 +2025-04-03 at 20:04:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ + +2025-04-03 at 20:04:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:04:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 mission location February 26 1966 +2025-04-03 at 20:04:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 20:04:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo launch and mission SC-017 or SC-020 or SC-020 first lunar mission +2025-04-03 at 20:04:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 20:04:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 launch mission +2025-04-03 at 20:04:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 20:04:07 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:04:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 launch site March 1966 +2025-04-03 at 20:04:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:04:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 Toroidal test +2025-04-03 at 20:04:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:04:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:04:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: AS-201 SC-009 Supercircular test location +2025-04-03 at 20:04:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-03 at 20:04:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:04:12 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: White Sands Missile Range AS-201 +2025-04-03 at 20:04:12 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ + +2025-04-03 at 20:04:12 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:04:14 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:04:14 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:04:14 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, True, False, False, True] +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_correctness:82 - Student lengths: [193, 259, 231, 365, 740, 186] +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [7, 7, 7, 7, 7, 7] +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_correctness:84 - Average student length: 329.00 +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 7.00 +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_correctness:86 - Length ratio: 47.00 +2025-04-03 at 20:04:14 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 20:04:14 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.567 Âą 0.357 +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 3.00 Âą 2.52 +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [1, 1, 5, 4, 0, 7] +2025-04-03 at 20:04:14 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928 + +10.3.2 Postrecovery Inspection +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +The Iwo Jima's position was established accurately using a satellite navigation system. A navigation fix was obtained at 1814 G.m.t., April 17, l970, and the position of the ship at spacecraft landing was dead-reckoned back to the time of landing and determined to be 2l degrees 34.7 minutes south latitude and 165 degrees 23.2 minutes west longitude. At landing a radar range of 8o00 yards and a visual bearing of 158.9 degrees east of north (true heading) were obtained from which the command module landing point was determined to be 2l degrees 38 minutes 24 seconds south latitude and 165 degrees 2l minutes 42 seconds west longitude. This position is judged to be accurate to within 5o0 yards. +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5 +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 20:04:14 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +cause for optics coupling display unit indications ofoptics movement during the Guidance and Navigation Performance check of zero optics mode operation Unable to perform tests on optical unit due to sslt water contanination 109018 zero optics mode To investigate the failure ofthe 0.0)g indication during entry Verifythe O.o5g entryzonitor system circuit,check the con- nectors,lamp,and wiring Not complete +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla. +------ +Result 2: +Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966 +------ + +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 5.0 +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.833 +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 20:04:14 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.233, Perfect scores: 1/6 +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.33 +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 20:04:14 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:04:14 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.469, Max reward: 0.872 +2025-04-03 at 20:04:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:04:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 'S-IVB lunar orbit duration +2025-04-03 at 20:04:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 20:04:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "S-IVB vibration duration Apollo 11" +2025-04-03 at 20:04:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:04:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: failures during first orbit on STS-1 +2025-04-03 at 20:04:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 20:04:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the duration of the S-IVB vibration on the Surveyor mission? +2025-04-03 at 20:04:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:04:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:04:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: launch +2025-04-03 at 20:04:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ + +2025-04-03 at 20:04:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "S-IVB vibration Apollo 11 translunar injection time" +2025-04-03 at 20:04:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:04:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration time during orbiter mission +2025-04-03 at 20:04:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:04:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the start and end hours of the S-IVB vibration on the Surveyor mission in 24-hour format? +2025-04-03 at 20:04:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:04:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:04:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Port Canaveral launch conditions +2025-04-03 at 20:04:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +The structural evaluation is based on guidance and control data, cabin pressure measurements, conmand module acceleration data, photographs īŧŒ and crew comments . + +Based on measured command module accelerations and on simulations using actual launch wind data, lumar module loads were within structural limits during laumch and translurar injection. Loads during docking and service propulsion and descent prcpulsion maneuvers were also within structural limits. + +Data telemetered during the oxygen tank incident indicate the presence of body bending oscillations in the docked spacecraft. The associated amplitudes, however, were of a very low level, and bending loads in the critical docking-tmnel area were well below design limits. + +6.2 ELECTRICAL POWER +------ + +2025-04-03 at 20:04:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 S-IVB vibration start time" +2025-04-03 at 20:04:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:04:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration hours during translunar injection +2025-04-03 at 20:04:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 20:04:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What are the start and end hours of the S-IVB vibration episode at engine 5 on the Apollo 12 mission? +2025-04-03 at 20:04:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:04:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:04:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Port Canaveral launch specific guidance and control data +2025-04-03 at 20:04:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 20:04:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo 11 S-II ignition 00:05:32" +2025-04-03 at 20:04:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 20:04:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB midcourse correction vibrations +2025-04-03 at 20:04:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 20:04:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the start hour of the S-IVB vibration at engine 5 on the Apollo 12 mission? +2025-04-03 at 20:04:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:04:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:04:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration during S-IVB coast phase +2025-04-03 at 20:04:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 20:04:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB engine cutoff time during translunar injection +2025-04-03 at 20:04:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 20:04:28 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 S-IVB vibration start time engine 5 +2025-04-03 at 20:04:28 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:04:28 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:04:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration characteristics +2025-04-03 at 20:04:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 20:04:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB vibration during first midcourse correction +2025-04-03 at 20:04:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 20:04:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB command level vibrations +2025-04-03 at 20:04:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:04:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:04:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB translunar injection +2025-04-03 at 20:04:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 20:04:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: when did the Apollo 8 S-IVB start to vibrate +2025-04-03 at 20:04:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:04:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:04:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: free-return trajectory S-IVB translunar injection +2025-04-03 at 20:04:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:04:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo S-IVB vibration Apollo 8 nominal time +2025-04-03 at 20:04:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:04:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:04:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: translunar injection S-IVB +2025-04-03 at 20:04:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 20:04:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:04:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: translunar injection +2025-04-03 at 20:04:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 20:04:37 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:04:37 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:04:37 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, True, False, False] +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_correctness:82 - Student lengths: [2028, 414, 323, 275, 347, 2024] +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [15, 15, 15, 15, 15, 15] +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_correctness:84 - Average student length: 901.83 +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 15.00 +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_correctness:86 - Length ratio: 60.12 +2025-04-03 at 20:04:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 20:04:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.608 Âą 0.433 +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 4.67 Âą 3.77 +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [10, 4, 0, 6, 0, 8] +2025-04-03 at 20:04:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Attempt to impact the expended S-IVB stage on the lunar surface within 350 km of the targeted impact point of 3 degrees south latitude and 30 degrees west longitude under nominal flight control conditions to excite the Apollo l2 seismometer. + +b. Postflight determination of the actual time and location of S-IVF impact to within. l second. + +Both objectives were accomplished, and the results are documented in reference 2. The impact was successfully detected by the seismometer and is reported in greater detail in section ll.3. + +Seven scientific experiments, in addition to those contained in the lunar surface experiment package, were also assigned as follows: + +a. Lunar field geology (S-059) b. Pilot describing function (T-029) c. Solar wind composition (S-080) d. S-band transponder exercise (S-164) e. Downlink bistatic radar observations of the moon (S-170) f. Gegenschein observation from lunar orbit (S-178) g。 Lunar surface closeup photography (S-184) +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +The structural evaluation is based on guidance and control data, cabin pressure measurements, conmand module acceleration data, photographs īŧŒ and crew comments . + +Based on measured command module accelerations and on simulations using actual launch wind data, lumar module loads were within structural limits during laumch and translurar injection. Loads during docking and service propulsion and descent prcpulsion maneuvers were also within structural limits. + +Data telemetered during the oxygen tank incident indicate the presence of body bending oscillations in the docked spacecraft. The associated amplitudes, however, were of a very low level, and bending loads in the critical docking-tmnel area were well below design limits. + +6.2 ELECTRICAL POWER +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Event Time, hr:min:sec Range zero - 19:13:00:00 G.m.t., Apri1 1l, 1970 Lift-off - 19:13:00.65 G.m.t., April 1l, 1970 S-IC outboard engine cutoff S-II engine igmition (command time) Launch escape tower jettison S-II engine cutoff S-IVB engine ignition (command time) S-IVB engine cutoff Translunar injection maneuver S-IVB/command and service module separation Docking Spacecraft ejection S-IVB separation maneuver First midcourse correction (service propulsion) Cryogenic oxygen tank incident Second midcourse correction (descent propulsion) S--IVB lunar impact Transearth injection (aescent propulsion) Third midcourse correction (descent propulsion) Fourth midcourse correction (LM reaction control) 00:02:44 00:02:45 00:03:21 00:09:53 00:09:54 00:12:30 02:35:46 03:06:39 03:19:09 04:01:01 04:18:01 30 :40 :50 55:54:53 61:29:43 77:56: 40 79:27:39 105:18:28 +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 20:04:37 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The Mission Control Center and the Manned Space Flight Network provided excellent support throughout this aborted mission. Minor problems occurred at different sites around the network, but all were corrected with no consequence to flight control support. Momentary data losses occurred seven different times as a result of power amplifier faults, computer processor executive buffer depletion, or wave guide faults. On each occasion, data lock-up was regained in just a few minutes. + +10.3 RECOVERY OPERATIONS +------ +Result 2: +At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry. +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ +Result 2: +Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57 + +(b) Transearth +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles. +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 20:04:37 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 20:04:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.267, Perfect scores: 1/6 +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 2.00 +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 20:04:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:04:37 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.356, Max reward: 0.772 +2025-04-03 at 20:04:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:04:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material is used for insulation between the shells of helium tank gas cylinders +2025-04-03 at 20:04:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 20:04:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: material used in helium tank insulation helium-3 containment +2025-04-03 at 20:04:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:04:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: heat shield helium tank insulation +2025-04-03 at 20:04:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:04:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material was used for the insulation between helium tank shells +2025-04-03 at 20:04:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 20:04:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material is used for insulation in a helium tank +2025-04-03 at 20:04:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:04:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material is used as insulation between the two shells of the helium tank for the Lunar surface gas storage. +2025-04-03 at 20:04:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:04:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:04:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Properties of aluminized Mylar insulation used in gas cylinders +2025-04-03 at 20:04:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 20:04:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank insulation materials +2025-04-03 at 20:04:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 20:04:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What material is used for the inner and outer shells of a helium tank +2025-04-03 at 20:04:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:04:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacecraft helium tank insulation material +2025-04-03 at 20:04:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:04:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:04:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Temperature stability of aluminized Mylar insulation used for helium tanks +2025-04-03 at 20:04:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 20:04:47 | INFO | src.agent:run_tool_calls:115 - īŋŊīŋŊīŋŊ Search Query: apollo helium tank inner shell insulation material +2025-04-03 at 20:04:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 20:04:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Materials used for helium tanks +2025-04-03 at 20:04:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 20:04:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacecraft helium tank insulation material stainless steel +2025-04-03 at 20:04:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:04:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:04:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Properties of aluminized Mylar insulation for high-temperature applications +2025-04-03 at 20:04:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:04:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure rise rate helium tank insulation material +2025-04-03 at 20:04:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:04:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vacuum jackets materials for helium tanks +2025-04-03 at 20:04:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:04:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacecraft helium tank insulation material +2025-04-03 at 20:04:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:04:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:04:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: vacuum jacket materials with low thermal conductivity +2025-04-03 at 20:04:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 20:04:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacecraft helium tank material +2025-04-03 at 20:04:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:04:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:04:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: materials for cryogenic heat exchanger tubes or jackets +2025-04-03 at 20:04:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:04:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Stainless steel helium tank insulation +2025-04-03 at 20:04:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 20:04:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:04:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: materials used for inner and outer shells of helium tank +2025-04-03 at 20:04:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:04:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacecraft helium tank material +2025-04-03 at 20:04:56 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:04:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:04:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: materials for inner and outer shells of cryogenic helium tank +2025-04-03 at 20:04:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:04:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:04:59 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:04:59 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:04:59 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, True, False, True, True] +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.67 +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_correctness:82 - Student lengths: [410, 358, 558, 233, 179, 312] +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [16, 16, 16, 16, 16, 16] +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_correctness:84 - Average student length: 341.67 +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 16.00 +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_correctness:86 - Length ratio: 21.35 +2025-04-03 at 20:04:59 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 20:04:59 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.700 Âą 0.255 +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 4.17 Âą 2.67 +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [4, 1, 4, 1, 8, 7] +2025-04-03 at 20:04:59 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:04:59 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe most likely cause of the anomaly is a tank-insulation degradation which w...'] +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:04:59 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nThe most likely cause of the anomaly is a tank-insulation degradation which w...', 'Result 1:\nThe most likely cause of the anomaly is a tank-insulation degradation which w...', "Result 1:\nhelium tanks, one at the manufacturer's plant and the other at the Manned Spa...", 'Result 1:\nFigure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure.\n...', 'Result 1:\nThe most likely cause of the anomaly is a tank-insulation degradation which w...', 'Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...', 'Result 1:\nFigure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure.\n...', 'Result 1:\nTemperature changes were noted in bays 3 and 4 of the service module in respo...'] +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 20:04:59 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.600, Perfect scores: 2/6 +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 6/6 +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 3.17 +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 2/6 +2025-04-03 at 20:04:59 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:04:59 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.326, Max reward: 0.647 +2025-04-03 at 20:05:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:05:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the pitch of a spacecraft? +2025-04-03 at 20:05:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 20:05:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft initial pitch angle +2025-04-03 at 20:05:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:05:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: initial pitch of spacecraft +2025-04-03 at 20:05:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 20:05:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:05:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the Apollo spacecraft's initial pitch configuration? +2025-04-03 at 20:05:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:05:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space shuttle initial pitch attitude +2025-04-03 at 20:05:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:05:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: common spacecraft initial pitch maneuvers +2025-04-03 at 20:05:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 20:05:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:05:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacecraft initial pitch configuration +2025-04-03 at 20:05:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:05:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space shuttle earth distant initial pitch attitude +2025-04-03 at 20:05:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 20:05:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pitch maneuver calculations for Apollo lunar mission +2025-04-03 at 20:05:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 20:05:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:05:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo command and lunar module initial pitch configuration +2025-04-03 at 20:05:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 20:05:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space shuttle initial pitch attitude before launch +2025-04-03 at 20:05:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-03 at 20:05:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft initial pitch maneuver +2025-04-03 at 20:05:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:05:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:05:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spaceships initial pitch configuration +2025-04-03 at 20:05:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:05:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space shuttle pitch angle launch +2025-04-03 at 20:05:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 20:05:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: launch pitch attitude and initial spacecraft pitch +2025-04-03 at 20:05:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 20:05:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:05:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Command Module pitch configuration +2025-04-03 at 20:05:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 20:05:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: manually controlled lunar module entry pitch adjustment +2025-04-03 at 20:05:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 20:05:15 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: launch pitch angle Saturn V +2025-04-03 at 20:05:15 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-03 at 20:05:15 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:05:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Command Module specifications +2025-04-03 at 20:05:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:05:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module pitchup maneuver +2025-04-03 at 20:05:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 20:05:18 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V launch pitch and electrostatic potentials +2025-04-03 at 20:05:18 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:05:18 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:05:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Command Module C-1 specifications +2025-04-03 at 20:05:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 20:05:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: manual pitch maneuver lunar module separation +2025-04-03 at 20:05:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 20:05:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V pitch angle during launch electrostatic potential +2025-04-03 at 20:05:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:05:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:05:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Command Module pitch after launch +2025-04-03 at 20:05:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 20:05:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: moon landing site and entry descent +2025-04-03 at 20:05:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 20:05:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V launch vehicle pitch angle and electrostatic potential +2025-04-03 at 20:05:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:05:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:05:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo Command Module pitch orientation after launch +2025-04-03 at 20:05:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Upon completion of photography, the two docked vehicles were maneuvered back to the service module separation attitude in preparation for the command module alignment. Star observation through the command module optics in this attitude was poor because of light reflecting from the lunar module, and the Conmander varied the pitch attitude by approximately 20 degrees in an attempt to improve star visibility. These attitude excursions, however, were not effective, and the spacecraft was returned to the original separation attitude for the command module alignment. + +8.9.5 Command Module Activation +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 20:05:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar landing and mission command module +2025-04-03 at 20:05:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 20:05:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Saturn V launch vehicle pitch and electrostatic potential +2025-04-03 at 20:05:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:05:25 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:05:25 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:05:25 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 1/6 answers correct +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1505, 1764, 1792, 526, 517, 365] +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [4, 4, 4, 4, 4, 4] +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_correctness:84 - Average student length: 1078.17 +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 4.00 +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_correctness:86 - Length ratio: 269.54 +2025-04-03 at 20:05:25 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 20:05:25 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.475 Âą 0.475 +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 5.00 Âą 5.00 +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [10, 10, 10, 0, 0, 0] +2025-04-03 at 20:05:25 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +DESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNAR MODULE ¡¡¡¡¡¡¡ A-1 A.3 EXPERIMENT EQUIPMENT ¡ A-2 A.4 LAUNCH VEHICLE ¡¡¡¡¡¡ A-5 A.5 MASS PROPERTIES .. A-5 Section Page APPENDIX B - SPACECRAFT HISTORIES B-1 APPENDIX C - POSTFLIGHT TESTING C-1 APPENDIX D - DATA AVAILABILITY D-1 APPENDIX E - MISSION REPORT SUPPLEMENTS E-1 REFEREN CES R-1 +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness. +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +Upon completion of photography, the two docked vehicles were maneuvered back to the service module separation attitude in preparation for the command module alignment. Star observation through the command module optics in this attitude was poor because of light reflecting from the lunar module, and the Conmander varied the pitch attitude by approximately 20 degrees in an attempt to improve star visibility. These attitude excursions, however, were not effective, and the spacecraft was returned to the original separation attitude for the command module alignment. + +8.9.5 Command Module Activation +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 20:05:25 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nThe unprecedented powered-down state of the command module required generatio...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nDESCRIPTIONS¡¡¡¡¡¡¡¡¡¡ A-1 A.1 COMMAND AND SERVICE MODULES .¡¡¡¡ A-1 A.2 LUNA...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nUpon completion of photography, the two docked vehicles were maneuvered back ...'] +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g. +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +At approximately 105 hours, the crew performed a manual descent propulsion maneuver to improve the entry angle. Since the primary guidance and navigation system was powered down, alignment was accomplished manually. The spacecraft was maneuvered to place the cusps of the earth' terminator on the Y-axis reticle of the crewmen optical alignment sight. The illuminated portion of the earth was then placed at the top of the reticle. This procedure positioned the lunar module X-axis perpendicular to the earth's terminator and permitted a retrograde maneuver to be performed perpendicular to the flight path to steepen the entry angle. The proper pitch attitude was maintained by positioning the sun in the top center portion of the telescope. With the spacecraft in the proper attitude, a body-axis alignment using the abort guidance system was followed immediately by entry into an attitude hold mode. This sequence resulted in attitude indications of zero for all axes and permitted use of the +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +Following the lunar module maneuver to the service module separation attitude, the comnand module platform heaters were activated, the command module reaction control system was pressurized, and each individual thruster was fired. An abort guidance attitude reference was provided with all zeros displayed on the attitude error needles. The lunar module was placed in an attitude hold mode using the abort guidance system; X-axis translation was monitored on the displays. After the reaction control system check was completed, the Conmander conducted a plus-X translation maneuver of 0.5 ft/sec, followed immediately by service module jettison. The pyro activation was heard and a minus 0.5-ft/sec translation maneuver was immediately commenced to remove the previously added velocity and preclude service module recontact. The jettison dynamics caused the undocked vehicles to pitch down about 10 degrees. Control was then switched to primary guidance minimum impulse, and a pitchup maneuver was +------ + +2025-04-03 at 20:05:25 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nAt lift-off, measured winds, both at the surface and in the region of maximum...', 'Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...', 'Result 1:\nAt approximately 105 hours, the crew performed a manual descent propulsion ma...', 'Result 1:\nFollowing the lunar module maneuver to the service module separation attitude...', 'Result 1:\nFollowing separation and translation, a manual pitch maneuver of 1.5 deg/sec ...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...'] +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll. +------ +Result 2: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +At lift-off, measured winds, both at the surface and in the region of maximum dynamic pressure, and accelerometer data indicate that structural loads were well below the established limits during all phases of flight. The predicted and calculated spacecraft loads at lift-off, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to or less than previous Apollo Saturn V launches. Command module accelerometer data prior to S-Ic center-engine cutoff indicate longitudinal oscillations similar to those measured on previous flights. Although longitudinal oscillations in the S-II engine structure and propellant system caused early shutdown of the center engine, the vibrations at the spacecraft during S-Il boost had an amplitude less than 0.05g at a frequency of l6 hertz. The maximum oscillation measured during either of the two S-IVB thrust periods was 0.o6g, also at a frequency of 16 hert z. Oscillations during all four launch vehicle +------ +Result 2: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more. + +Supplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation +------ +Result 2: +The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V. + +ll.l.2 Very-Low and Low-Frequency Radio Noise +------ + +2025-04-03 at 20:05:25 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAll attitude control functions were satisfactory. Initial separation from the...', 'Result 1:\nAt lift-off, measured winds, both at the surface and in the region of maximum...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...', 'Result 1:\nIt is known that the electrostatic potentials develop on jet aircraft. These ...'] +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-03 at 20:05:25 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-03 at 20:05:25 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1 +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll + +Velocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed. +2025-04-03 at 20:05:25 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 20:05:25 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 20:05:25 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:05:25 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.114, Max reward: 0.436 +2025-04-03 at 20:05:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:05:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is left-hand electrical circuit interrupter +2025-04-03 at 20:05:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 20:05:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +What gets jammed or stuck in a left-hand electrical circuit interrupter? +2025-04-03 at 20:05:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 20:05:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What type of energy or material is often associated with electrical circuit interrupters and is considered a fire hazard if not approached carefully? +2025-04-03 at 20:05:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-03 at 20:05:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: arc fault circuit interrupter left-hand side components +2025-04-03 at 20:05:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 20:05:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is an electrical arc that is 'escaped' from a left-hand circuit interrupter +2025-04-03 at 20:05:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-03 at 20:05:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:05:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: left-hand electrical circuit interrupter in command module propulsion system +2025-04-03 at 20:05:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 20:05:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Incident reports of gas or liquid escaping from an electrical circuit interrupter in a lunar module. +2025-04-03 at 20:05:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-03 at 20:05:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: failures of left-hand command module service module umbilical guillotine electric arc fault circuit interrupter +2025-04-03 at 20:05:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 20:05:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:05:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Material escaped from lunar module electrical circuit interrupter +2025-04-03 at 20:05:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 20:05:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module service module umbilical guillotine left-hand circuit failure +2025-04-03 at 20:05:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 20:05:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:05:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what material escaped from lunar module electrical circuit when jettisoning +2025-04-03 at 20:05:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 20:05:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module service module umbilical guillotine propellant gas leakage cause +2025-04-03 at 20:05:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ + +2025-04-03 at 20:05:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:05:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What materials can escape from the automated transfer valve during a lunar spacecraft operation +2025-04-03 at 20:05:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 20:05:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: service propulsion auxiliary propellant gaging system failure +2025-04-03 at 20:05:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 20:05:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:05:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What type of material escaped from the lunar module as a result of the reaction control system +2025-04-03 at 20:05:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 20:05:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: oxygen tank propulsive venting maneuver command module +2025-04-03 at 20:05:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ + +2025-04-03 at 20:05:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:05:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission: material escaped from reaction control valves +2025-04-03 at 20:05:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ +Result 2: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ + +2025-04-03 at 20:05:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection thermal control mode difficulties command module +2025-04-03 at 20:05:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ + +2025-04-03 at 20:05:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:05:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What materials make up the components of the Apollo lunar module's reaction control valves +2025-04-03 at 20:05:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 20:05:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command module translation hand controller issues +2025-04-03 at 20:05:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 20:05:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:05:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo spacecraft: metallic materials used in the lunar module reaction control system +2025-04-03 at 20:05:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 20:05:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: labrador translation maneuver issues +2025-04-03 at 20:05:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ + +2025-04-03 at 20:05:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:05:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module reaction control system components +2025-04-03 at 20:05:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 20:05:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: index of l containing commands and spacecraft dynamics during oxygen tank anomaly +2025-04-03 at 20:05:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ +Result 2: +All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +------ + +2025-04-03 at 20:05:48 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:05:48 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:05:48 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, False, True] +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_correctness:82 - Student lengths: [537, 125, 591, 863, 1245, 349] +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [14, 14, 14, 14, 14, 14] +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_correctness:84 - Average student length: 618.33 +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 14.00 +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_correctness:86 - Length ratio: 44.17 +2025-04-03 at 20:05:48 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 20:05:48 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.517 Âą 0.341 +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 4.00 Âą 4.28 +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [2, 1, 0, 10, 10, 1] +2025-04-03 at 20:05:48 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 20:05:48 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +At 2-l/2 hours prior to entry, the command module was fully powered up and lunar module power transfer was terminated. After command module computer activation, the unfavorable spacecraft attitude delayed communications signal lockup and the ensuing ground uplink commands. The stable platform was coarse aligned to ground-supplied reference angles, and an optical fine alignment made using two stars. Particles venting from the command module umibilical area impeded command module optics operation. With the lunar module attached to the command moaule and the command module optics pointed away from the sun, individual stars were barely visible through the optics. Also sun reflections from the lumar module sublimator and the nearest reaction control quad prevented positive identification of constellations. + +8.9.6 Lunar Moaule Undocking +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ +Result 2: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +The experiment consisted of an electronics package and two sensor probes which were to be placed in bore holes, predrilled by the crew using the Apollo lunar surface drill. At each end of the probe was a gradient heat sensor with heater coil, a ring sensor lo centimeters from each end, and four thermocouples in the probe cable. The probe consisted of two $55$ -centimeter sections joined at a 2-inch spacing with a flexible spring. + + + +Figure A-2.- Experiment subpackage number 2. + +A.3.2 Charged Particle Lunar Environment Experiment +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover. + +A.4 LAUNCH VEHICLE + +Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation. + +A.5 MASS PROPERTIES +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970 +------ +Result 2: +and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969 +------ + +2025-04-03 at 20:05:48 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nThe next series of events occurred within a fraction of a second between the ...', 'Result 1:\nAll mechanical systems functioned properly. One mechanical anomaly, however, ...', 'Result 1:\nAll mechanical systems functioned properly. One mechanical anomaly, however, ...', 'Result 1:\nAll mechanical systems functioned properly. One mechanical anomaly, however, ...', 'Result 1:\nDuring periods when the lunar module and the command module cabins were inter...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nThe auxiliary dump nozzle was used for the first time on an Apollo mission. D...', 'Result 1:\nDuring periods when the lunar module and the command module cabins were inter...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...', 'Result 1:\nPerformance of Lunar Module Reaction Control System Review 6 Ascent Propulsio...'] +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The two interrupters open the electrical circuits about 30 milliseconds before the wires are severed by the command module/service module umbilical guillotine. As illustrated in the figure, a cam fork is moved by a piston, which is operated by propellant gas from redundant cartridges, to function a lift plate. Motion of this plate disconnects the male and female portions of electrical connectors located, respectively, in the lift plate and in the base plate of the interrupter. At the completion of the stroke, the fork is brought to rest by impacting and crushing an aluminum block mounted on the interrupter housing. +------ +Result 2: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Service module.- At the time the system was powered down, reaction control system propellant usage was l08 poumds higher than predicted. The higher usage is attributed to the increased thruster activity required to null the effects of propulsive venting from both oxygen tanks during the incident. The usages listed in the following table were calculated from telemetered helium tank pressure data using the relationship between pressure, volume, and temperature. + +Fuel, 1b Oxi di zerīŧŒ lb Loaded Quad A Quad B Quad( C Quad D 110.4 109.5 110.1 110.1 440.1 225.6 225.5 225.4 226.2 902.7 336.0 335.0 335.5 336.3 1342.8 Consumed Remaining at time 286* 1056.8 + +Preflight planned usage was 178 pounds. +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +The command module reaction control system helium pressures and temperatures and the helium manifold pressures were normal from lift-off to system activation just prior to entry. The pressures before activation reflected the general cooling of the system resulting from the powered dowr configuration of the command module. The helium source temperatures dropped from $70^{\circ}$ toabout $35^{\circ}$ F during the mission. Prior to system activation the lowest engine injector temperature was $1.5^{\circ}$ F. A preheat cycle brougnt injector temperatures to acceptable levels and hot firing checks were satisfactory . + +Just prior to undocking, two injector temperatures were $5^{\circ}$ Fbelow minimum. However, engine operation was expected to be normal, despite the low temperatures, and undocking was performed without heating the engines. +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +A.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Platform-sensed velocity changes, ft/sec Command module axes Lunar module axes X Z X Y Z Service module separation PlusX translation Minus X translation ed up at separation Platformnotpower- 0.67 -1.90 -0.08 0.01 0.01 +0'0- + +Table 6.4-I summarizes the pertinent control system parameters during each translation maneuver. Spacecraft dynamic response during all maneuvers was normal.. + +The throttle profile for the first midcourse correction performed by the lunar module was 5 seconds at 12.7 percent followed by 27 seconds at 40 percent. The firing was preceded by a l0-second, four-jet ullage maneuver. A number of plus-X firings occurred during the maneuver because pitch and roll thrusters were not inhibited by a Verb 65 entry, as required by the checklist. +------ +Result 2: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly . + +The only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table. +------ +Result 2: +All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission. +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3. +------ +Result 2: +Based upon an analysis of the interrupter design, its location, and its relationship to adjacent equipment, it is concluded that gas will not escape prior to the completion of the deadfacing function and that, should such escape occur, the gas will not adversely affect any other components. Therefore, no hardware modification is necessary. + +This anomaly is closed. + +14.2 LUNAR MODULE + +14.2.l Abnormal Supercritical Helium Pressure Rise +------ + +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 20:05:48 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.133, Perfect scores: 0/6 +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 2/6 +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.50 +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 20:05:48 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:05:48 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.404, Max reward: 0.749 +2025-04-03 at 20:05:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:05:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: rise-rate of helium tank pressure during Titan II rocket countdown demonstration test +2025-04-03 at 20:05:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:05:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure range helium tank rise-rate characteristics Apollo 1 countdown demonstration test +2025-04-03 at 20:05:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:05:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise-rate pressure range during countdown demonstration test +2025-04-03 at 20:05:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:05:53 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: near-meltdown helium tank rise rate CN-0759 ST465 SCTR-46 +2025-04-03 at 20:05:53 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:05:53 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:05:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pressure range for helium tank rise rate increase during Titan II countdown demonstration test +2025-04-03 at 20:05:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:05:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: abnormal helium pressure rise rate above 900 psia Apollo 1 countdown demonstration test +2025-04-03 at 20:05:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:05:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo helium tank pressure range and rise rate +2025-04-03 at 20:05:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:05:55 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (640-900) helium tank pressure rise rate test +2025-04-03 at 20:05:55 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 20:05:55 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:05:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: upper bound of pressure range above 900 psia where helium tank rise rate increased during Titan II countdown demonstration test +2025-04-03 at 20:05:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:05:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo helium tank rise rate pressure range lower bound +2025-04-03 at 20:05:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:05:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: supercritical helium tank pressure increase 900 psia +2025-04-03 at 20:05:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 20:05:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:05:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lower bound pressure for helium tank rise rate +2025-04-03 at 20:05:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:05:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise rate 900-1000 psia +2025-04-03 at 20:05:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:05:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:06:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lower bound pressure helium tank rise rate abnormal increase +2025-04-03 at 20:06:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-03 at 20:06:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 937 psia rise rate helium tank +2025-04-03 at 20:06:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:06:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise rate critical pressure +2025-04-03 at 20:06:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 20:06:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: 937 psia helium tank helium temperature rise rate +2025-04-03 at 20:06:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:06:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank pressure rise rate range bounds +2025-04-03 at 20:06:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar helium tank pressure rise rate 640-900 +2025-04-03 at 20:06:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:06:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise rate pressure upper bound +2025-04-03 at 20:06:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank pressure 900 psia +2025-04-03 at 20:06:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 20:06:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:06:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise rate pressure bounds +2025-04-03 at 20:06:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Helium tank rise rate 937 psia after first descent engine firing +2025-04-03 at 20:06:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 20:06:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:06:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise rate upper bound +2025-04-03 at 20:06:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:09 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:06:09 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:06:09 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, True, False, False, False] +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_correctness:82 - Student lengths: [368, 581, 180, 443, 1683, 1847] +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [24, 24, 24, 24, 24, 24] +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_correctness:84 - Average student length: 850.33 +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 24.00 +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_correctness:86 - Length ratio: 35.43 +2025-04-03 at 20:06:09 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-03 at 20:06:09 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.508 Âą 0.393 +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 4.00 Âą 4.04 +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 3, 2, 0, 10, 9] +2025-04-03 at 20:06:09 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-03 at 20:06:09 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-03 at 20:06:09 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates. + +The average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 20:06:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:06:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 20:06:10 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 20:06:10 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:06:10 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:06:10 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 20:06:10 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 20:06:10 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 20:06:10 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:06:10 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.267, Perfect scores: 1/6 +2025-04-03 at 20:06:10 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 20:06:10 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.50 +2025-04-03 at 20:06:10 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 20:06:10 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:06:10 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.349, Max reward: 0.709 +2025-04-03 at 20:06:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:06:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: combine throttle position sensor failure and throttle position with formula one +2025-04-03 at 20:06:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ + +2025-04-03 at 20:06:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the maximum throttle position of a typical commercial airliner or general aviation aircraft? +2025-04-03 at 20:06:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 20:06:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the final throttle position before landing in aircraft +2025-04-03 at 20:06:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:06:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: mechanical throttle position +2025-04-03 at 20:06:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 20:06:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: throttle position final aircraft +2025-04-03 at 20:06:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 20:06:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:06:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Formula One throttle position malfunction and attitude error needles not nulled +2025-04-03 at 20:06:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 20:06:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth maneuver apollo spacecraft throttle profile +2025-04-03 at 20:06:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 20:06:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what was the final throttle position during transearth injection maneuver +2025-04-03 at 20:06:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:06:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft throttle position transearth injection final +2025-04-03 at 20:06:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:06:21 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: throttle final position space shuttle +2025-04-03 at 20:06:21 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:06:21 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:06:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Formula One throttle position attitude error needles not nulled acceleration drift +2025-04-03 at 20:06:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ + +2025-04-03 at 20:06:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection throttle 10 percent +2025-04-03 at 20:06:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:06:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection final throttle position +2025-04-03 at 20:06:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:06:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: relation 12.6% throttle to full throttle +2025-04-03 at 20:06:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 20:06:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:06:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Revs throttle position not nulled after midcourse correction +2025-04-03 at 20:06:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 20:06:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo descent propulsion throttle minimal blowdown mode +2025-04-03 at 20:06:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 20:06:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection final descent throttle +2025-04-03 at 20:06:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 20:06:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth throttle profile analogy +2025-04-03 at 20:06:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ + +2025-04-03 at 20:06:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:06:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Thorpe roll manual switch engine firing +2025-04-03 at 20:06:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 20:06:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection cardin rocket throttle +2025-04-03 at 20:06:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:06:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection throttle profile +2025-04-03 at 20:06:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:06:34 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: primary guidance system throttle profile +2025-04-03 at 20:06:34 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 20:06:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:06:37 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: acceleration mode engine firing throttle +2025-04-03 at 20:06:37 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 20:06:37 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:06:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Thurber throttle position after midcourse correction +2025-04-03 at 20:06:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 20:06:38 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:06:38 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:06:38 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, True, False, False, True] +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_correctness:82 - Student lengths: [2037, 390, 451, 1553, 712, 366] +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [13, 13, 13, 13, 13, 13] +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_correctness:84 - Average student length: 918.17 +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 13.00 +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_correctness:86 - Length ratio: 70.63 +2025-04-03 at 20:06:38 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 20:06:38 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.637 Âą 0.352 +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 5.50 Âą 4.43 +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [7, 5, 5, 14, 0, 2] +2025-04-03 at 20:06:38 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ + +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ + +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 20:06:38 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nCondition Maneuver Second midcourse correction Transearth injection Third mid...', 'Result 1:\nAfter completing the maneuver to the attitude required for the final midcours...', 'Result 1:\nUncompens ated Error term error One-sigma specification Offset velocity, ft/s...', 'Result 1:\nThe passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not ...', 'Result 1:\nThe passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not ...', 'Result 1:\nThe passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not ...', 'Result 1:\nCondition Maneuver Second midcourse correction Transearth injection Third mid...'] +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Low cabin temperature, resulting from a greatly reduced thermal loading from powered down electrical equipment, was uncomfortable to the crew during the return flight. For most of this time, power levels were maintained between 350 and 400 watts. Environnental equipment operation, however, was normal for this thermal loading, with temperatures of the Water/glycol coolant at the sublimator inlet of approximately $46^{\circ}\texttt{F}$ Cabin temperatures were typically between $54^{\circ}$ and $60^{\circ}$ F, and suit inlet temperatures were maintained between $40^{\circ}$ and 41° F during this portion of the flight. +------ + +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 20:06:38 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:06:39 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...'] +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:06:39 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...'] +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ + +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# äēē +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44 Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 Âą0.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60 +------ +Result 2: +Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth +------ + +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ + +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +------ +Result 2: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ + +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-03 at 20:06:39 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver. + +8.9 TRANSEARTH COAST + +8.9.1 Coast Phase Activities +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:06:39 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe descent propulsion system performed normally during the 34.3- second midc...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...'] +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 1.0 +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.167 +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 20:06:39 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.567, Perfect scores: 2/6 +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 4/6 +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 2.00 +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 3/6 +2025-04-03 at 20:06:39 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:06:39 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.404, Max reward: 0.637 +2025-04-03 at 20:06:40 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:06:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot training documents aviation certification +2025-04-03 at 20:06:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 20:06:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot training section aircraft manual +2025-04-03 at 20:06:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ + +2025-04-03 at 20:06:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilots training manual +2025-04-03 at 20:06:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-03 at 20:06:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: flight training manual pilot instruction section +2025-04-03 at 20:06:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ + +2025-04-03 at 20:06:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot training procedures +2025-04-03 at 20:06:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ +Result 2: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ + +2025-04-03 at 20:06:44 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: piloting training manual +2025-04-03 at 20:06:44 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-03 at 20:06:44 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:06:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot training aviation field trips extravehicular equipment +2025-04-03 at 20:06:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 20:06:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot training emergency procedures aircraft operation +2025-04-03 at 20:06:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 20:06:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Pilot Command Module Pilot Instrumentation Occupational Training +2025-04-03 at 20:06:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:06:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot training manual astronaut instruction section +2025-04-03 at 20:06:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-03 at 20:06:46 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo pilot training philippines +2025-04-03 at 20:06:46 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 20:06:46 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:06:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: extravehicular equipment space exploration modifications '''extravehicular mobility unit'' +2025-04-03 at 20:06:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-03 at 20:06:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: pilot training manual section aircraft operations +2025-04-03 at 20:06:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ + +2025-04-03 at 20:06:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 Pilot Command Module Pilot Occupational Training geology +2025-04-03 at 20:06:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:06:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 training lunar geology +2025-04-03 at 20:06:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-03 at 20:06:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:06:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo training astronaut extravehicular equipment modifications +2025-04-03 at 20:06:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 20:06:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 training manual +2025-04-03 at 20:06:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-03 at 20:06:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 lunar surface training objectives +2025-04-03 at 20:06:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-03 at 20:06:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:06:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo command module modifications life support systems emergency procedures +2025-04-03 at 20:06:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 20:06:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 8 pilots training section +2025-04-03 at 20:06:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 20:06:52 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:06:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 mission report medical personnel physiology astronauts +2025-04-03 at 20:06:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-03 at 20:06:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew training schedule and geology training +2025-04-03 at 20:06:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:06:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:06:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 heart rate stress physiology +2025-04-03 at 20:06:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 20:06:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 mission operations report section PIIOTS +2025-04-03 at 20:06:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:06:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:06:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 crew physiology stress training +2025-04-03 at 20:06:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 20:06:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission crew briefings or pilot training sections +2025-04-03 at 20:06:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ + +2025-04-03 at 20:06:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:07:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 12 crew training stress management +2025-04-03 at 20:07:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 20:07:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 incident report noise vibration +2025-04-03 at 20:07:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 20:07:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:07:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 13 crew lunar landing training geology training +2025-04-03 at 20:07:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 20:07:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 impact and seismic signals +2025-04-03 at 20:07:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 20:07:04 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:07:04 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:07:04 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, True, False, True] +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1875, 1728, 428, 269, 546, 312] +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [3, 3, 3, 3, 3, 3] +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_correctness:84 - Average student length: 859.67 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 3.00 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_correctness:86 - Length ratio: 286.56 +2025-04-03 at 20:07:04 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_format:228 - Responses ending properly: 5/6 +2025-04-03 at 20:07:04 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.700 Âą 0.224 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 5.00 Âą 3.65 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [10, 10, 3, 2, 1, 4] +2025-04-03 at 20:07:04 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ +Result 2: +The structure in the forward end of the docking tunnel was reinforced to accommodate the expected higher parachute loads due to the increased weight of the command module. In the sequential system the timing signal which disables the roll engines during service module separation was changed from a 5.5- to a 2-second interval, and a cutoff time of 25 seconds was incorporated for the translation engines instead of allowing them to fire until the propellant was depleted. These timing changes were instituted to minimize the effects of fuel slosh and to improve service-module separation characteristics. The stripline units in the high-gain antenna were changed to an improved design. A detachable filter was provided for installing over the cabin heat exchanger exhaust to assist in collection of free lumar dust after crew transfer from the lunar module. An extra urine filter, in addition to the primary and backup units, was stowed and could be used to reduce the possibility of a clogged +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The thickness of the outer-skin shielding for the forward hatch was increased from 0.004 to 0.010 inch to improve the resistance to the tearing that was noted on Apollo l2. The D-ring handle on the modularized equipment storage assembly was changed to a looped cable to simplify the deployment operation. The thermal insulation for the landing gear was modified to reduce the total insulation weight by 27.2 pounds. Both a color and a black-and-white television camera were included for increased reliability of television coverage on .the lunar surface. The primary guidance programs were modified to permit reentry into the automatic and attitude hold modes of operation after manual control was exercised; this change was incorporated to provide improved final descent capability in the event of obscuration from lunar dust. The event timer was modified so that after it counted down to zero, it would count up automatically and thus reduce crew workload during critical events. The descent +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ +Result 2: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications. + +Prior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows. + +Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12 +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload. + +During the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry. + +9.2 INFLIGHT HISTORY + +9.2.l Adaptation to Weightlessness +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes +------ + +2025-04-03 at 20:07:04 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nuse of field radios, extravehicular equipment, and assistance from mission co...', 'Result 1:\nuse of field radios, extravehicular equipment, and assistance from mission co...', 'Result 1:\nThe extravehicular mobility unit underwent several modifications to improve l...', 'Result 1:\nThe thickness of the outer-skin shielding for the forward hatch was increased...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nThe biomedical data were excellent in quality during the period from launch t...', 'Result 1:\nAt 55:54:54, a telemetry dropout was observed. Immediately after the incident...', 'Result 1:\nAt 55:54:54, a telemetry dropout was observed. Immediately after the incident...', 'Result 1:\nAt 55:54:54, a telemetry dropout was observed. Immediately after the incident...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ +Result 2: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ +Result 2: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan. + +This anomaly is closed. + +14.l.3 Shaft Fluctuations in the Zero Optics Mode + +Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission. +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +An unexplained characteristic of the S-IVB impact is the rapid buildup from its beginning to the peak value. This initial stage of the signal cannot be explained solely by the scattering of seismic waves in a rubbletype material, as was thought possible from the ascent stage impact data. Several alternate hypotheses are under study, but no firm conclusions have been reached. Signal scattering, however, may explain the character of the later part of the signal. + +The fact that such precise targeting accuracy was possible for the S-IVB impact, with the resulting seismic signals so large, have greatly encouraged seismologists to study possible future S-IvB impacts. For ranges extended to 500 kilometers, the data return could provide a means for determining moon structures to depths approaching 200 kilometers. + +The four primary objectives (see ref. 7) assigned to the Apollo 13 mission were as follows: +------ +Result 2: +Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds. +------ + +2025-04-03 at 20:07:04 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\ndiscussion of the flight and associated crew activities is presented in secti...', 'Result 1:\nuse of field radios, extravehicular equipment, and assistance from mission co...', 'Result 1:\ndiscussion of the flight and associated crew activities is presented in secti...', 'Result 1:\nSupplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reco...', 'Result 1:\nCommander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., an...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCommander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., an...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nTo guard against operational problems of this type in the future, a caution n...', 'Result 1:\nAn unexplained characteristic of the S-IVB impact is the rapid buildup from i...'] +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:07:04 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nThe performance of the abort guidance system and all attitude control aspects...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +discussion of the flight and associated crew activities is presented in section 8.0 as a complementary description to this section. +------ +Result 2: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +PREPARED BY + +Mission Evaluation Team + +APPROVED BY + + + +NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970 + + + +TABLE OF CONTENTS + +Section Page +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-03 at 20:07:04 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\ndiscussion of the flight and associated crew activities is presented in secti...', 'Result 1:\nPREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND...'] +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +use of field radios, extravehicular equipment, and assistance from mission control during these field trips made the training more effective. +------ +Result 2: +The television presentation during the midcourse correction maneuver, as well as during transposition and docking, interfered with normal operational functions to a degree not seen in training. The lunar module pilot was forced to spend full time adjusting, pointing, and narrating the television broadcast. A suggested alternative for telecasting during dynamic events is to have the ground do all commentary. Crew-designated television can be conveniently performed during a lull period when full attention can be given to presentation requirements. + +8.7.3 Cryogenic Oxygen Tank Incident + +At approximately 55 hours 54 minutes, a loud noise was heard when the Command Module Pilot was in the left seat, the Commander in the lower equipment bay, and the Lunar Module Pilot in the tunnel. The noise was comparable to that noted in exercising the lunar module repressurization valve. The Command Module Pilot and Lunar Module Pilot also reported a minor vibration or tremor in the spacecraft. +------ + +2025-04-03 at 20:07:04 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nuse of field radios, extravehicular equipment, and assistance from mission co...'] +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: 1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The performance of the abort guidance system and all attitude control aspects of the digital autopilot were nominal. Following the service module oxygen tank anomaly, power was applied to the primary guidance system for use in establishing passive thermal control modes and to maintain attitude control until the transearth injection maneuver. + +The passive thermal control mode after transearth injection was initiated using the digital autopilot in the manual minimum impulse mode. The crew had considerable difficulty in establishing acceptable initial conditions for the passive thermal control mode. This difficulty was largely caused by the necessity to use the translation hand controller +------ +Result 2: +Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr. + +8.0 PIIOTS' REPORT + +8.1 TRAINING +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as "observers" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive +------ +Result 2: +a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites. + +Thirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned: + +TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS +------ + +2025-04-03 at 20:07:04 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nThe performance of the abort guidance system and all attitude control aspects...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...', 'Result 1:\nCrew training for Apollo 13 commenced on August l, 1969. The crew was based i...'] +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 20:07:04 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.233, Perfect scores: 0/6 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.83 +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 20:07:04 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:07:04 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.485, Max reward: 0.646 +2025-04-03 at 20:07:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:07:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Is there a specific type of testing (e.g. is it electrical, mechanical, pressure testing ) that validates isolation valves in space exploration systems? +2025-04-03 at 20:07:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 20:07:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "spacecraft isolation valve testing protocols" +2025-04-03 at 20:07:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 20:07:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"spacecraft isolation valve electrical safety testing methods" +2025-04-03 at 20:07:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 20:07:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: predictive analytics for spacecraft electrical testing +2025-04-03 at 20:07:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 20:07:09 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: leakage current testing spacecraft isolation valves +2025-04-03 at 20:07:09 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 20:07:09 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:07:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What type of resistance testing is used to verify isolation valve wiring in spacecraft? +2025-04-03 at 20:07:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 20:07:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "spacecraft isolation valve wiring testing verification" +2025-04-03 at 20:07:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 20:07:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"spacecraft electrical wiring resistance testing methods" +2025-04-03 at 20:07:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:07:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: validation of potable water quantity fluctuation isolation valve resistance +2025-04-03 at 20:07:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +Potable water was obtained by periodically pressurizing the potable tank with surge-tank oxygen and withdrawing potable water until the pressures equalized. This method provided potable water for crew use until 24 hours prior to entry, at which time water could not be withdrawn from the potable tank and it appeared to be exhausted [section 5.8]. + +The hatch, probe, and drogue were secured in the couches by lap belt and shoulder harness restraints to prevent movement during subsequent maneuvers. + +8.7.4 Midcourse Correction to a Free Return +------ + +2025-04-03 at 20:07:13 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space shuttle isolation valve testing procedure +2025-04-03 at 20:07:13 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 20:07:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:07:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What type of electrical test is capable of revealing a miswiring issue through an inductive coupling interaction? +2025-04-03 at 20:07:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-03 at 20:07:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Spacecraft isolation valve isolation verification testing" +2025-04-03 at 20:07:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 20:07:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"apollo spacecraft isolation valve testing resistance measurement" +2025-04-03 at 20:07:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 20:07:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft potable water tank fill and emptying methods +2025-04-03 at 20:07:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 20:07:16 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: resistance testing isolation valve wiring assurance +2025-04-03 at 20:07:16 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:07:16 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:07:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Isolation valve testing requirements" +2025-04-03 at 20:07:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 20:07:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"aluminum vs electroless nickel plating in space electronics" +2025-04-03 at 20:07:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 20:07:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: potable water leakage and freezing point under low temperatures +2025-04-03 at 20:07:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 20:07:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: integrated test for isolation valve wiring +2025-04-03 at 20:07:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 20:07:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:07:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Spacecraft isolation valve testing procedures" +2025-04-03 at 20:07:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 20:07:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"apollo 13 isolation valve isolation wiring" +2025-04-03 at 20:07:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ + +2025-04-03 at 20:07:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: gas bubbles in potable water tank lead to water freezing point +2025-04-03 at 20:07:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:07:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: isolation valve electrical connection validation +2025-04-03 at 20:07:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 20:07:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:07:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Isolation valve testing protocols for space applications" +2025-04-03 at 20:07:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 20:07:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +" proced of testing isolation valves for wiring shorts" +2025-04-03 at 20:07:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:07:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank contamination and thermal conductivity in space +2025-04-03 at 20:07:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 20:07:26 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: isolation valve voltage dependency testing +2025-04-03 at 20:07:26 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:07:26 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:07:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: " Methods for verifying isolation valve wiring" +2025-04-03 at 20:07:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 20:07:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"ap9 isolation valve electrical fault testing" +2025-04-03 at 20:07:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 20:07:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank contamination anomalies and pressure rise rate testing +2025-04-03 at 20:07:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:07:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: isolation valve voltage tolerance testing +2025-04-03 at 20:07:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 20:07:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:07:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Isolation valve validation requirements" +2025-04-03 at 20:07:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:07:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"validation of wire connections isolation valves through inductive coupling" +2025-04-03 at 20:07:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 20:07:32 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: hydrogen contamination and tank insulation in supercritical helium tanks +2025-04-03 at 20:07:32 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:07:32 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:07:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Isolation valve qualification for spacecraft systems" +2025-04-03 at 20:07:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 20:07:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: +"isolated wiring electronica failure mode test" +2025-04-03 at 20:07:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-03 at 20:07:35 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: hydrogen contamination and thermal conductivity effects in vacuum insulated tanks +2025-04-03 at 20:07:35 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 20:07:35 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:07:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: plumbing component failure and heater switch operation +2025-04-03 at 20:07:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 20:07:36 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:07:36 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:07:37 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 6/6 answers correct +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, True] +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_correctness:82 - Student lengths: [920, 1798, 1034, 1908, 2058, 597] +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [17, 17, 17, 17, 17, 17] +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_correctness:84 - Average student length: 1385.83 +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 17.00 +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_correctness:86 - Length ratio: 81.52 +2025-04-03 at 20:07:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 20:07:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.663 Âą 0.347 +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 6.50 Âą 3.77 +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [3, 10, 0, 9, 10, 7] +2025-04-03 at 20:07:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-03 at 20:07:37 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: [] +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +During periods when the lunar module and the command module cabins were interconnected, the lunar module and command module cabin pressure readings were approximately equal, verifying the operation of the command module cabin pressure transducers. + + + +(c) 142:45 through 142:56 hours. Figure l4-l2.- Concluded. + +The suit measurement indicated correctly during the brief instrumentation power-up periods at l02 and 123 hours. However, just prior to entry, the suit indication was approximately 0.3 psi lower than cabin pressure but increased to 7.7 psia when the cabin pressure was reading 13.9 psia just prior to landing. + +This transducer also behaved erratically on Apollo l2. Postflight analysis of both the Apollo l2 and Apollo l3 transducers determined the cause to be internal contamination from electroless nickel plating particles. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +MSC-02680 + +DISTRIBUTION AND REFERENCING + +This paper is not suitable for general distribution or referencing. It may be referenced only in other working correspondence and documents by participating organizations. + +MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER1970 +------ +Result 2: +1.0 SUMMARY 1-1 2.0 INTRODUCTION¡. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... ¡¡¡ 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .¡ 5-1 5.2 ELECTRICAL POWER ¡¡¡¡ 5-2 5.3 CRYOGENIC STORAGE.¡¡¡ 5-3 5.4 COMMUNICATIONS EQUIPMENT ¡ 5-4 5.5 INSTRUMENTATION.¡¡¡¡¡¡¡ 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .¡ 5-5 5.7 REACTION CONTROL.¡¡¡¡¡¡¡ 5-11 5.8 ENVIRONMENTAL CONTROL .¡. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL ¡¡¡ 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .¡ 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION ¡¡¡ 6-8 6.7 ENVIRONMENTAL CONTROL.¡¡¡ 6-9 7.0 MISSION CONSUMABLES ¡¡¡¡¡. ÂˇÂˇã€Âˇ 7-1 7.1 COMMAND AND SERVICE MODULES .¡¡¡¡ 7-1 7.2 LUNAR MODULE ¡¡¡¡¡ 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION ’· 8-2 8.6 TRANSPOSITION AND DOCKING .¡.. 8-7 +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +electronic box and trigger the antenna logic to produce the scan-limit functions when the voltage sum reaches a threshold value. Under normal operating conditions, the threshold voltage is reached when the C-axis angular travel is between 95 and 1l5 degrees. + +The failure mode of the C-axis induction potentiometer is a short in the stator excitation winding. Shorting one half of the stator's primary winding to ground would produce a greater slope in the curye showing the induction potentiometer transformation ratio versus angular travel. This slope increase would produce nonlinear effects because the magnetic flux is concentrated in one-half of the primary winding. Further analysis is in progress to establish the particular failure and what might have caused the condition. + +A test will be performed at the launch site on future spacecraft to preclude launching with either a bad C-axis or A-axis generator. + +An anomaly report will be published when the analysis is complete. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +Potable water was obtained by periodically pressurizing the potable tank with surge-tank oxygen and withdrawing potable water until the pressures equalized. This method provided potable water for crew use until 24 hours prior to entry, at which time water could not be withdrawn from the potable tank and it appeared to be exhausted [section 5.8]. + +The hatch, probe, and drogue were secured in the couches by lap belt and shoulder harness restraints to prevent movement during subsequent maneuvers. + +8.7.4 Midcourse Correction to a Free Return +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure. + +A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity. + +This anomaly is closed. + +14.2.2 Abnormal Descent Stage Noise +------ +Result 2: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +After initial cryogenic oxygen filling during the countdown demonstration test at Kennedy Space Center, tank 2 could not be detanked using the normal procedures. The problem resulted from loose or misaligned plumbing components in the dog-leg portion of the tank fill path. After numerous attempts using gaseous oxygen purges and higher expulsion pressures, the fluid was boiled off through the use of the tank heaters and fans, assisted by pressure cycling. During the detanking sequence, the heaters were on for about 8 hours, but it was believed that no damage would be sustained by the tank or its components because of the protection afforded by intermal thermal switches. However, the use of the heaters in detanking required that the switches open under a load of 6 amperes at 65 V dc, twice the normal flight operating conditions, for each heater. Tests show that opening the switches under these conditions will fuse the contacts closed and eventually damage fan motor wire insulation. It +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition. + +This anomaly is closed. + + + +Figure. l4-l0.- Tunnel gusset protection. + +14.l.7 Reaction Control Isolation Valve Failure + +During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ +Result 2: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The miswired valve (fig. 14-ll) passed the functional checks during buildup and checkout because, even with the closing coil lead completely disconnected, the valve can be closed through an inductive coupling with the oxidizer-valve closing coil. That is, a reverse-polarity voltage can be generated in the oxidizer valve opening coil through a "transformer" + + + +Figure l4-ll.- Isolation valve circuit. + +action. This voltage is applied to the fuel valve opening coil where it induces a magnetic field flux that closes the fuel valve. With 28 volts or more on the spacecraft bus, this phenomenon was consistently repeatable. With 24 to 28 volts on the bus, the valve would occasionally close, and with less than 24 volts, the valve would not close. Since preflight testing is accomplished at 28 volts, the functional tests did not disclose the miswiring. During the mission, the voltage was such that the valve did not close when commanded and therefore was found open after the fli ght. +------ +Result 2: +Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired. + +This anomaly is closed. + +14.l.8 Potable Water Quantity Fluctuations +------ + +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 5.0 +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.833 +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.373 +2025-04-03 at 20:07:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 20:07:37 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:07:37 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.299, Max reward: 0.658 +2025-04-03 at 20:07:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:07:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Where was the original location of the sensing point for the water separator drain tank located in a rollover protection system, in a truck or ship? +2025-04-03 at 20:07:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 20:07:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: boats water separator drainage location origin +2025-04-03 at 20:07:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +During the periods when it was activated, the command module environmental control system performed normally. From the time of powering dowm at approximately 58 hours until reactivation approximately l-1/2 hours before entry, environmental control for the interconnected cabins was maintained using lumar module equipment. Two anomalies associated with the environmental control instrumentation occurred and are discussed in sections 14.l.8 and l4.l.9. An additional discrepancy, noted after landing and discussed in section l0.3, was the position of the inlet postlanding ventilation valve at the time of recovery. This discrepancy is discussed in section 14.l.2. +------ + +2025-04-03 at 20:07:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "original location of sensing point for water separator drain tank" +2025-04-03 at 20:07:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-03 at 20:07:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the original location of the water separation tank sensing point in modern engines? +2025-04-03 at 20:07:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +accumulator in the center-engine liquid oxygen line is being incorporated on future vehicles to decouple the line from the crossbeam, and therefore suppress any vibration amplitudes. Addition of a vibration detection system which would monitor structural response in the l4-to-20 Hz range and initiate engine cutoff if vibrations approach a dangerous level. is also under investigation as a backup. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 20:07:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the location of water separator drain sensor in diesel fuel tank. +2025-04-03 at 20:07:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 20:07:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "nuclear submarine water treatment drains US Navy" +2025-04-03 at 20:07:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +After the command module auxiliary urine dump, used through the side hatch, was exercised, the crew was requested by the ground to inhibit all further overboard dumps so as not to interfere with navigation sightings. This single dump was noted to seriously degrade visibility through the command module hatch window. Since this restriction was never retracted, all subsequent urine collections were stowed onboard. The containers utilized for urine collections were the six lunar module urine transfer bags, three command module backup waste bags, the condensate container, two water collection bags for the portable life support system, and three urine collection devices. The command module waste stowage compartment appeared to be full with only seven fecal bags stowed in this area. Adding to the waste stowage problem was the stiffness of the outer fecal bags. +------ + +2025-04-03 at 20:07:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:07:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the original location of a sensing point for a water separator drain tank or similar component in a vehicle or system designed for fuel monitoring? +2025-04-03 at 20:07:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-03 at 20:07:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: submarine pumps leak location +2025-04-03 at 20:07:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ + +2025-04-03 at 20:07:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "water separator drain tank location" +2025-04-03 at 20:07:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 20:07:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the original location of the water separator drain tank sensing point in the pressure vessel used before redesign? +2025-04-03 at 20:07:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-03 at 20:07:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: space shuttle fuel cell water separator tank sensor location +2025-04-03 at 20:07:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:07:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo urine collectionHouston mission report" +2025-04-03 at 20:07:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the command module auxiliary urine dump, used through the side hatch, was exercised, the crew was requested by the ground to inhibit all further overboard dumps so as not to interfere with navigation sightings. This single dump was noted to seriously degrade visibility through the command module hatch window. Since this restriction was never retracted, all subsequent urine collections were stowed onboard. The containers utilized for urine collections were the six lunar module urine transfer bags, three command module backup waste bags, the condensate container, two water collection bags for the portable life support system, and three urine collection devices. The command module waste stowage compartment appeared to be full with only seven fecal bags stowed in this area. Adding to the waste stowage problem was the stiffness of the outer fecal bags. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:07:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:07:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Location of the water separator drain tank in a spacecraft or vehicle +2025-04-03 at 20:07:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 20:07:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: focus 14 1l oxygen manifold +2025-04-03 at 20:07:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +accumulator in the center-engine liquid oxygen line is being incorporated on future vehicles to decouple the line from the crossbeam, and therefore suppress any vibration amplitudes. Addition of a vibration detection system which would monitor structural response in the l4-to-20 Hz range and initiate engine cutoff if vibrations approach a dangerous level. is also under investigation as a backup. +------ + +2025-04-03 at 20:07:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "surfraft water separator drain tank location" +2025-04-03 at 20:07:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 20:07:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of water separator drain sensing point on MoonEUSS spacesuit +2025-04-03 at 20:07:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 20:07:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fuel cell water separator tank pressure sensor +2025-04-03 at 20:07:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:07:49 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo mission urine storage compartment" +2025-04-03 at 20:07:49 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the command module auxiliary urine dump, used through the side hatch, was exercised, the crew was requested by the ground to inhibit all further overboard dumps so as not to interfere with navigation sightings. This single dump was noted to seriously degrade visibility through the command module hatch window. Since this restriction was never retracted, all subsequent urine collections were stowed onboard. The containers utilized for urine collections were the six lunar module urine transfer bags, three command module backup waste bags, the condensate container, two water collection bags for the portable life support system, and three urine collection devices. The command module waste stowage compartment appeared to be full with only seven fecal bags stowed in this area. Adding to the waste stowage problem was the stiffness of the outer fecal bags. +------ +Result 2: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ + +2025-04-03 at 20:07:49 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:07:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Location of the water/glycol separator drain tank in the command and service module of a lunar spacecraft +2025-04-03 at 20:07:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 20:07:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: oxygen manifold ascent stage oxygen tank interface +2025-04-03 at 20:07:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ + +2025-04-03 at 20:07:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "surge tank water separator drain location" +2025-04-03 at 20:07:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:07:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of the carbon dioxide sensor versus water separator drain tank +2025-04-03 at 20:07:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 20:07:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of fuel cell water separator drain sensor +2025-04-03 at 20:07:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-03 at 20:07:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo mission wastewater separator location" +2025-04-03 at 20:07:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 20:07:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:07:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Definition and location of the water separator drain tank in the lunar landing module +2025-04-03 at 20:07:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 20:07:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: water separator drain tank oil or gas separator location +2025-04-03 at 20:07:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 20:07:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "ASHUR system water separator drain testing protocol" +2025-04-03 at 20:07:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-03 at 20:07:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of carbon dioxide sensor versus water separator drain tank in EUSS +2025-04-03 at 20:07:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 20:07:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: cryogenic oxygen tank 2 fuel cell oxygen shutoff valve pressure sensor +2025-04-03 at 20:07:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:07:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo wastewater tank bypass valve location" +2025-04-03 at 20:07:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 20:07:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:07:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: definition of Water separator drain tank +2025-04-03 at 20:07:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 20:07:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: command and service modules water separator drain tank location +2025-04-03 at 20:07:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-03 at 20:07:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "potable water tank drain location and instrumentation" +2025-04-03 at 20:07:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:07:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: water separator drain tank sensing point original location in Apollo missions +2025-04-03 at 20:07:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 20:07:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of fuel cell oxygen shutoff valve water separator drain sensor +2025-04-03 at 20:07:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-03 at 20:07:57 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "urine valve water separator drain tank location" +2025-04-03 at 20:07:57 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the command module auxiliary urine dump, used through the side hatch, was exercised, the crew was requested by the ground to inhibit all further overboard dumps so as not to interfere with navigation sightings. This single dump was noted to seriously degrade visibility through the command module hatch window. Since this restriction was never retracted, all subsequent urine collections were stowed onboard. The containers utilized for urine collections were the six lunar module urine transfer bags, three command module backup waste bags, the condensate container, two water collection bags for the portable life support system, and three urine collection devices. The command module waste stowage compartment appeared to be full with only seven fecal bags stowed in this area. Adding to the waste stowage problem was the stiffness of the outer fecal bags. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 20:07:57 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:08:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is a water separator drain tank +2025-04-03 at 20:08:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 20:08:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module reaction control system water separator drain tank +2025-04-03 at 20:08:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 20:08:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "potable water transducer location" +2025-04-03 at 20:08:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ + +2025-04-03 at 20:08:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: original sensing point water separator drain lunar module Apollo +2025-04-03 at 20:08:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 20:08:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: no information +2025-04-03 at 20:08:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ + +2025-04-03 at 20:08:00 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "urine collection system fluid valve location" +2025-04-03 at 20:08:00 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ +Result 2: +After the command module auxiliary urine dump, used through the side hatch, was exercised, the crew was requested by the ground to inhibit all further overboard dumps so as not to interfere with navigation sightings. This single dump was noted to seriously degrade visibility through the command module hatch window. Since this restriction was never retracted, all subsequent urine collections were stowed onboard. The containers utilized for urine collections were the six lunar module urine transfer bags, three command module backup waste bags, the condensate container, two water collection bags for the portable life support system, and three urine collection devices. The command module waste stowage compartment appeared to be full with only seven fecal bags stowed in this area. Adding to the waste stowage problem was the stiffness of the outer fecal bags. +------ + +2025-04-03 at 20:08:00 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:08:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How is a water separator drain tank used in a spacecraft or vehicle +2025-04-03 at 20:08:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 20:08:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module lithium hydroxide cartridges water separator drain tank +2025-04-03 at 20:08:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 20:08:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "transducer location relative to potable water tank" +2025-04-03 at 20:08:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ + +2025-04-03 at 20:08:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: SEAL System Water Separator drain tank Locator +2025-04-03 at 20:08:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 20:08:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: original Apollo mission location of water separator drain sensor +2025-04-03 at 20:08:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 20:08:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Apollo descent stage tank valve location" +2025-04-03 at 20:08:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 20:08:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:08:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: function of water separator drain tank in spacecraft +2025-04-03 at 20:08:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 20:08:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module command module water separator drain tank location +2025-04-03 at 20:08:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-03 at 20:08:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "potable water tank transducer" +2025-04-03 at 20:08:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:08:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: location of water separator drain tank in Apollo +2025-04-03 at 20:08:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:08:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo water separator drain sensor location upstream of suit fans +2025-04-03 at 20:08:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-03 at 20:08:06 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "oxygen manifold ascent stage valve location" +2025-04-03 at 20:08:06 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 20:08:06 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:08:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module water gun water separator drain tank +2025-04-03 at 20:08:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 20:08:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: water separator drain tank location Apollo spacecraft +2025-04-03 at 20:08:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 20:08:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: engine or suit water separator drain sensor location after suit modifications +2025-04-03 at 20:08:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:08:08 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "ascent stage oxygen manifold manipulator port valve" +2025-04-03 at 20:08:08 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 20:08:09 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:08:09 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:08:09 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 0/6 answers correct +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1971, 1936, 1806, 2017, 2059, 1587] +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [25, 25, 25, 25, 25, 25] +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_correctness:84 - Average student length: 1896.00 +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 25.00 +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_correctness:86 - Length ratio: 75.84 +2025-04-03 at 20:08:09 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_format:228 - Responses ending properly: 6/6 +2025-04-03 at 20:08:09 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.871 Âą 0.177 +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 10.00 Âą 1.00 +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [9, 10, 9, 10, 10, 12] +2025-04-03 at 20:08:09 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected. + +5.6 GUIDANCE, NAVIGATION, AND CONTROL +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +freezing in the lines most likely occurred. This freezing condition could have existed at the time a sample was to be taken. When the spacecraft was returned to the manufacturer's plant, 2l.3 pounds were drained from the potable tank. The water system was subsequently checked and was found to operate properly. Both the hot and cold potable water contained gas bubbles. To eliminate these gas bubbles, which had also been experienced on previous missions, a gas separator cartridge was provided but not used. +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ +Result 2: +During the periods when it was activated, the command module environmental control system performed normally. From the time of powering dowm at approximately 58 hours until reactivation approximately l-1/2 hours before entry, environmental control for the interconnected cabins was maintained using lumar module equipment. Two anomalies associated with the environmental control instrumentation occurred and are discussed in sections 14.l.8 and l4.l.9. An additional discrepancy, noted after landing and discussed in section l0.3, was the position of the inlet postlanding ventilation valve at the time of recovery. This discrepancy is discussed in section 14.l.2. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +accumulator in the center-engine liquid oxygen line is being incorporated on future vehicles to decouple the line from the crossbeam, and therefore suppress any vibration amplitudes. Addition of a vibration detection system which would monitor structural response in the l4-to-20 Hz range and initiate engine cutoff if vibrations approach a dangerous level. is also under investigation as a backup. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ +Result 2: +Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs. + +7.l.l Service Propulsion Propellants + +The service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off. + +Fuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines "79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7 + +7.1.2 Reaction Control Propellants +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +It became apparent that there were insufficient lithium hydroxide cartridges in the lumar module to support the abort mission, even with allowable'carbon dioxide levels extended to 8 partial pressure of l5 mm Hg. With ground instructions, a system was constructed which attached a command module lithium hydroxide cartridge to each of two lunar module suit hoses. The Commander's remaining hose was placed in the tunnel area to provide fresh oxygen to the command module, while the Lunar Module Pilot's remaining hose was positioned in the lunar module environmental control area. At a later time, a second cartridge was added in series to the cartridges initially installed, as shown in figure 6.7-l. In each case, the drop in carbon dioxide levels reported by the ground showed Satisfactory operation of this improvised carbon dioxide removal system.: +------ +Result 2: +advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ +Result 2: +The service module was separated $\scriptstyle4-3/4$ hours before entry , affording the crew an opportunity to observe and photograph the damaged bay 4 area. The command moduie was separated from the service module by using the lunar module reaction control system. The lunar module was retained for as long as possible to provide maximum electrical power in the command module for entry . +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +In using the lunar module water gun to dampen a towel, a piece of towel material most likely became caught in the gun nozzle when the actuating trigger was released, resulting in water leakage from the nozzle. The lunar module water gun was returned to earth and during postflight testing was found to be operating properly. Postflight testing also showed that reactuation of the valve can flush any towel material from the gun. The command module water gun was satisfactorily used for the remainder of the mission. + +7.0 MISSION_CONSUMABLES + +Consumables from the command and servi ce modules were used normally during the 56 hours prior to the incident, at a modified usage schedule for 2 hours after the incident, and after command module activation just prior to entry. The lumar module usages occurred in the period following power-up until the two spacecraft were undocked. + +7.1 COMMAND AND SERVICE MODULES +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 20:08:09 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\n(fig. 14-20) caused the leakage. When the valve is installed, the forward O-r...', 'Result 1:\nDuring the flight, the pressure in the ascent stage oxygen tank 2 increased, ...', 'Result 1:\nDuring the flight, the pressure in the ascent stage oxygen tank 2 increased, ...', 'Result 1:\nDuring the flight, the pressure in the ascent stage oxygen tank 2 increased, ...', 'Result 1:\nASHUR Purpose Tests performed Results Environmental Control 109007 To determi...', 'Result 1:\nThe service module was separated $\\scriptstyle4-3/4$ hours before entry , aff...', 'Result 1:\n6.7 ENVIRONMENTAL CONTROL\n\nEnvironmental. control system performance was sati...', 'Result 1:\nIt became apparent that there were insufficient lithium hydroxide cartridges ...', 'Result 1:\nIn using the lunar module water gun to dampen a towel, a piece of towel mater...', 'Result 1:\nIn using the lunar module water gun to dampen a towel, a piece of towel mater...'] +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:08:09 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nEfforts to install the tunnel hatch were terminated when the Commander observ...', 'Result 1:\nEfforts to install the tunnel hatch were terminated when the Commander observ...', 'Result 1:\nEfforts to install the tunnel hatch were terminated when the Commander observ...', 'Result 1:\nThe oxygen distribution system operated nominally until deactivation followin...', 'Result 1:\nASHUR Purpose Tests performed Results Environmental Control 109007 To determi...', 'Result 1:\nThe measurement is not essential for flight safety or mission success. The po...', 'Result 1:\nThe measurement is not essential for flight safety or mission success. The po...', 'Result 1:\nThe measurement is not essential for flight safety or mission success. The po...', 'Result 1:\nThe measurement is not essential for flight safety or mission success. The po...'] +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +accumulator in the center-engine liquid oxygen line is being incorporated on future vehicles to decouple the line from the crossbeam, and therefore suppress any vibration amplitudes. Addition of a vibration detection system which would monitor structural response in the l4-to-20 Hz range and initiate engine cutoff if vibrations approach a dangerous level. is also under investigation as a backup. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +The extravehicular mobility unit underwent several modifications to improve lunar surface capability. Scuff patches were added to the pressure garment assembly to prevent wear of the thermal/meteoroid garment caused by chaffing of the lunar boots. A device was added in the neck area of the pressure suit to provide drinking water to the crewmen during extravehicular activity. A center eyeshade was installed at the top of the extravehicular visor assembly to reduce incoming glare and to aid in dark adaptation when entering shadow. Abrasion cover gloves were included to be used over the extravehicular gloves to reduce wear and heat conduction during core drilling operations. The electrical connnector on the remote control unit for the portable life support system was redesigned to permit easier engagement. The manufacturing technique for the regulator in the oxygen purge system was modified to minimize the possibility of gas leakage. + +A.3 EXPERIMENT EQUIPMENT +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The only anomaly observed in the environmental control system was a reverse leakage from the oxygen manifold through the shutoff valve into the ascent oxygen tank 2. Following the use of oxygen from the tank on two occasions, tank pressure was permitted to increase to the regulated manifold pressure, where it remained for the duration of the flight. The maximum leakage rate through the valve was approximately 0.22 lb/hr. Both the specification leakage rate and the preflight test leakage rate were 0.0ol lb/hr. The leaking valve would have presented a problem if this ascent oxygen tank had developed an external leak. Further information regarding this anomaly is contained in section 14.2.4. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. + + + +Figure 6.7-l.- Supplemental carbon dioxide removal system. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ +Result 2: +The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0 + +7.1.4 Oxy ge n + +Following the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry. + +7.1.5 Water + +At the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation. + +7.1.6 Batteries +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The measurement is not essential for flight safety or mission success. The potable water tank is continually refilled with fuel cell product water, and when the potable water tank is full, fuel cell product water is automatically diverted to the waste water tank, which is periodically dumped overboard. Water from the potable water tank is used mainly for drinking and food reconstitution. Since fuel cell water generation rates can be computed from power generation levels and since potable water usage rates can be estimated with reasonable accuracy, the quantity of water in the potable water tank can be determined with acceptable accuracy without the quantity measurement. + +This anoma.ly is closed. + +14.l.9 Suit Pressure Transducer Failure +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ +Result 2: +Unusual variations in the oxygen flow rates to all three fuel cells were observed in the 3-minute period preceding the tank pressure loss. These variations were caused by the simultaneous pressure excursions taking place in cryogenic oxygen tank 2. The fuel cell l regulated nitrogen pressure indication went to the lower limit of the measurement when the pressure in cryogenic oxygen tank 2 dropped. Analysis of related fuel cell parameters confirmed this discrepancy to be a loss of instrumentation readout and not an actual loss of the regulated nitrogen pressure. Performance of fuel cells l and 3 degraded within 3 minutes after the Oxygen tank 2 pressure dropped. The degradation is considered to have been caused by the fuel cell oxygen shutoff valves closing abruptly because of the shock generated when the bay 4 panel separated. A more detailed discussion is contained in reference l. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Table E-I contains a listing of all supplemental reports that are or will be published for the Apollo 7 through Apollo l3 mission reports. Also indicated in the table is the present status of each report not published or the publication date for those which have been completed. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +This transducer has operated erratically on two previous missions. Testing after Apollo 8 traced the failure during that mission to moisture contamination within the transducer. Similar fluctuations noted during Apollo l2 were traced to a minute quantity of undetermined contamination on the surface of the resistance wafer. Characteristically, the signal level decreased first to indicate an increase in the resistance but returned to more normal readings as the wafer cleaned itself. Disassembly of the Apollo l3 transducer and water tank did not produce evidence of either contamination or corrosion. The spacecraft wiring which could have produced the problem was checked and no intermittents were found. +------ +Result 2: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +Inspection also revealed that both the cabin and suit loop pressure transducers contained various contaminants identified as solder flux, glass beads (0.o4 mm diameter), and fibers from the wipers used in the transducer manufacturer's clean room; all of which could potentially affect the transducer operation. + +To assure that one of the pressure transducers is operative, the Apollo 14 cabin pressure transducer will be disassembled, the plating will be inspected and the instrument will be cleaned, reassembled and installed. + +For Apollo 15 and subsequent, the suit and cabin pressure transducers will be disassembled and cleaned. The plating will be inspected for cracking or flaking and the units will be reassembled. The suit pressure transducers will be reassembled without soldering. + +This anomaly is closed. + +14.l.l0 Gas Leak in Electrical Circuit Interrupter +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +------ +Result 2: +ASHUR Purpose Tests performed Results Environmental Control 109007 To determine contaninates present or damage incurred in 9o0 psi system Anelyze the oxygen filters upstream o restrictors and check valves for contaminates. Perform acceptance test of oxygen ir:air regulator 109008 To determine contaninates present in residual oxygen in surge tank snd repressurization package Withdraw sample and analyze for contaminates No rigrificant difference from the araiysis per- formed at:adine 109016 To investigate the failure of the postlanding ventilation valve to cycle open Determine positionofinletvaive mechanical safety pin.Attenpt to operate valve,ther renove for failure analysis Not complete 109020 Todetermine the cause of failure othe suit pressure transaucer Perform calibration check,dis- assembly,and failure anaysis Not compiete 109021 Todetermine the cause of failure o!the potable water transducer Remove,disassemble,and per- form failure analysis Hot complete 109015 To investigate the +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again. +------ +Result 2: +After the command module auxiliary urine dump, used through the side hatch, was exercised, the crew was requested by the ground to inhibit all further overboard dumps so as not to interfere with navigation sightings. This single dump was noted to seriously degrade visibility through the command module hatch window. Since this restriction was never retracted, all subsequent urine collections were stowed onboard. The containers utilized for urine collections were the six lunar module urine transfer bags, three command module backup waste bags, the condensate container, two water collection bags for the portable life support system, and three urine collection devices. The command module waste stowage compartment appeared to be full with only seven fecal bags stowed in this area. Adding to the waste stowage problem was the stiffness of the outer fecal bags. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +After the command module auxiliary urine dump, used through the side hatch, was exercised, the crew was requested by the ground to inhibit all further overboard dumps so as not to interfere with navigation sightings. This single dump was noted to seriously degrade visibility through the command module hatch window. Since this restriction was never retracted, all subsequent urine collections were stowed onboard. The containers utilized for urine collections were the six lunar module urine transfer bags, three command module backup waste bags, the condensate container, two water collection bags for the portable life support system, and three urine collection devices. The command module waste stowage compartment appeared to be full with only seven fecal bags stowed in this area. Adding to the waste stowage problem was the stiffness of the outer fecal bags. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +After the command module auxiliary urine dump, used through the side hatch, was exercised, the crew was requested by the ground to inhibit all further overboard dumps so as not to interfere with navigation sightings. This single dump was noted to seriously degrade visibility through the command module hatch window. Since this restriction was never retracted, all subsequent urine collections were stowed onboard. The containers utilized for urine collections were the six lunar module urine transfer bags, three command module backup waste bags, the condensate container, two water collection bags for the portable life support system, and three urine collection devices. The command module waste stowage compartment appeared to be full with only seven fecal bags stowed in this area. Adding to the waste stowage problem was the stiffness of the outer fecal bags. +------ +Result 2: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +After the command module auxiliary urine dump, used through the side hatch, was exercised, the crew was requested by the ground to inhibit all further overboard dumps so as not to interfere with navigation sightings. This single dump was noted to seriously degrade visibility through the command module hatch window. Since this restriction was never retracted, all subsequent urine collections were stowed onboard. The containers utilized for urine collections were the six lunar module urine transfer bags, three command module backup waste bags, the condensate container, two water collection bags for the portable life support system, and three urine collection devices. The command module waste stowage compartment appeared to be full with only seven fecal bags stowed in this area. Adding to the waste stowage problem was the stiffness of the outer fecal bags. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ +Result 2: +After the command module auxiliary urine dump, used through the side hatch, was exercised, the crew was requested by the ground to inhibit all further overboard dumps so as not to interfere with navigation sightings. This single dump was noted to seriously degrade visibility through the command module hatch window. Since this restriction was never retracted, all subsequent urine collections were stowed onboard. The containers utilized for urine collections were the six lunar module urine transfer bags, three command module backup waste bags, the condensate container, two water collection bags for the portable life support system, and three urine collection devices. The command module waste stowage compartment appeared to be full with only seven fecal bags stowed in this area. Adding to the waste stowage problem was the stiffness of the outer fecal bags. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The auxiliary dump nozzle was used for the first time on an Apollo mission. Dumping through this nozzle was discontinued and urine was subsequently stored onboard because a considerable number of particles were evident on the hatch window and these interfered with navigation sightings。. + +Upon recovery, the outlet valve of the postlanding ventilation was open and the inlet valve was closed, whereas both valves should have been open. This condition is reported in section l0.3.2, and the anomaly is discussed in section 14.l.2. + +The performance of the lumar module systems is discussed in this section. All systems that are not discussed either performed as intended Or were not used. Discrepancies and aromalies are generally mentioned but are discussed in greater detail in the Anomaly Summary, sections 14.2 and14.3. + +6.1 STRUCTURAL +------ +Result 2: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction. + +The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring + + + +Figure 14-l9.- Oxygen-supply system. +------ +Result 2: +(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination. + +The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly. + + + +Figure l4-20.- Ascent stage tank shutoff valve:. +------ + +2025-04-03 at 20:08:09 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nPreflight testing of both command module and lunar module water supplies reve...', 'Result 1:\nAfter the command module auxiliary urine dump, used through the side hatch, w...', 'Result 1:\nAfter the command module auxiliary urine dump, used through the side hatch, w...', 'Result 1:\nThe command module was powered up with the three entry batteries, which had b...', 'Result 1:\nThe auxiliary dump nozzle was used for the first time on an Apollo mission. D...', 'Result 1:\nAfter the command module auxiliary urine dump, used through the side hatch, w...', 'Result 1:\nThe auxiliary dump nozzle was used for the first time on an Apollo mission. D...', 'Result 1:\nThe auxiliary dump nozzle was used for the first time on an Apollo mission. D...', 'Result 1:\nDuring the flight, the pressure in the ascent stage oxygen tank 2 increased, ...', 'Result 1:\nDuring the flight, the pressure in the ascent stage oxygen tank 2 increased, ...'] +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 3.0 +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.500 +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.500 +2025-04-03 at 20:08:09 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.300, Perfect scores: 0/6 +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 3.33 +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 20:08:09 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:08:09 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.332, Max reward: 0.624 +2025-04-03 at 20:08:11 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:08:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank rise rate prelaunch +2025-04-03 at 20:08:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:08:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the normal prelaunch-standby rise rate of a heel tone helium tank for a typical aircraft or military aircraft? +2025-04-03 at 20:08:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:08:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the standard helium tank pressure rise rate during the descent phase of a launch or depressurization of a space shuttle or helium buoys? +2025-04-03 at 20:08:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 20:08:14 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: typical helium tank prelaunch standby rise rate +2025-04-03 at 20:08:14 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:08:14 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:08:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank normal prelaunch rise rate NASA standard +2025-04-03 at 20:08:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:08:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the expected prelaunch-standby rise rate for a standard helium tank for the Apollo spacecraft? +2025-04-03 at 20:08:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:08:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the standard prelaunch-standby rise rate of a helium tank used in space shuttle launches? +2025-04-03 at 20:08:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:08:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:08:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: helium tank pressure rise rate definition NASA +2025-04-03 at 20:08:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 20:08:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What was the standard prelaunch-standby rise rate for a helium tank in the Apollo spacecraft? +2025-04-03 at 20:08:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:08:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: What is the helium tank rise rate at 640 psia? +2025-04-03 at 20:08:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:08:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:08:22 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: (iterate) What is the helium tank rise rate during the heat-leak test, which results in a rise rate of 7.9 psi/hour at 640 psia? +2025-04-03 at 20:08:22 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 20:08:22 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:08:23 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:08:23 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:08:23 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 4/6 answers correct +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, True, True, False, False, False] +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.50 +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.50 +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_correctness:82 - Student lengths: [136, 120, 298, 334, 1013, 357] +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [12, 12, 12, 12, 12, 12] +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_correctness:84 - Average student length: 376.33 +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 12.00 +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_correctness:86 - Length ratio: 31.36 +2025-04-03 at 20:08:23 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 20:08:23 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.408 Âą 0.318 +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 1.83 Âą 1.57 +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [3, 3, 4, 1, 0, 0] +2025-04-03 at 20:08:23 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ + +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +------ +Result 2: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ + +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 20:08:23 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: [] +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard. + +The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process. +2025-04-03 at 20:08:23 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 4.0 +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.667 +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.471 +2025-04-03 at 20:08:23 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.400, Perfect scores: 2/6 +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.83 +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 2/6 +2025-04-03 at 20:08:23 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:08:23 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.342, Max reward: 0.775 +2025-04-03 at 20:08:24 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:08:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: anomalies data patterns +2025-04-03 at 20:08:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ + +2025-04-03 at 20:08:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: anomalies in data definition +2025-04-03 at 20:08:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ + +2025-04-03 at 20:08:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Specificity of financial market anomalies +2025-04-03 at 20:08:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ + +2025-04-03 at 20:08:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: anomalies in economics +2025-04-03 at 20:08:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +(section ll.3). +------ + +2025-04-03 at 20:08:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: medical anomalies locations +2025-04-03 at 20:08:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ +Result 2: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-03 at 20:08:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:08:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo anomaly data inspection command module +2025-04-03 at 20:08:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 20:08:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 15 anomaly command module +2025-04-03 at 20:08:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 20:08:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Earth observations anomalies finance vs space exploration +2025-04-03 at 20:08:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 20:08:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: gyrocompassing misalignment error coefficient velocity data +2025-04-03 at 20:08:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 20:08:30 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: launch site 6 laser anomalies +2025-04-03 at 20:08:30 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ + +2025-04-03 at 20:08:30 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:08:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module command module separation tunnel hatch +2025-04-03 at 20:08:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 20:08:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 15 anomaly investigation +2025-04-03 at 20:08:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:08:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spac NASA navigation optical measurements +2025-04-03 at 20:08:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Upon completion of photography, the two docked vehicles were maneuvered back to the service module separation attitude in preparation for the command module alignment. Star observation through the command module optics in this attitude was poor because of light reflecting from the lunar module, and the Conmander varied the pitch attitude by approximately 20 degrees in an attempt to improve star visibility. These attitude excursions, however, were not effective, and the spacecraft was returned to the original separation attitude for the command module alignment. + +8.9.5 Command Module Activation +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 20:08:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: gyrocompassing misalignment null bias drift inertial measurement unit +2025-04-03 at 20:08:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 20:08:33 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electromagnetic anomalies during launch field test +2025-04-03 at 20:08:33 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 20:08:33 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:08:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module re-entry separation procedures +2025-04-03 at 20:08:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 20:08:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: apollo 15 cryogenic oxygen supply anomaly +2025-04-03 at 20:08:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:08:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo mission translunar navigation +2025-04-03 at 20:08:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 20:08:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: gyrocompassing misalignment null bias drift S-IvB instrument unit platform alignment +2025-04-03 at 20:08:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 20:08:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electromagnetic field differences site 4 vs sites 1 2 3 +2025-04-03 at 20:08:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 20:08:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 5 prompts +2025-04-03 at 20:08:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: reaction control system lunar module reentry procedures +2025-04-03 at 20:08:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 20:08:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo translunar manual navigation +2025-04-03 at 20:08:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 20:08:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: spacecraft gyrocompassing misalignment null bias drift null-bias drift gyroscope alignment +2025-04-03 at 20:08:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 20:08:39 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electromagnetic emissions launch vehicle +2025-04-03 at 20:08:39 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ +Result 2: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ + +2025-04-03 at 20:08:39 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:08:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module reentry descent routine +2025-04-03 at 20:08:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ + +2025-04-03 at 20:08:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo translunar separation and docking procedures +2025-04-03 at 20:08:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:08:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module platform alignment null bias drift comparative inertial navigation systems +2025-04-03 at 20:08:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 20:08:42 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: electric field comparison site 4 vs sites l 2 3 7 +2025-04-03 at 20:08:42 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 20:08:42 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:08:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module emergency descent system +2025-04-03 at 20:08:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ + +2025-04-03 at 20:08:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB translunar injection attitude and velocity +2025-04-03 at 20:08:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 20:08:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection maneuver lunar module spacecraft navigation accuracy +2025-04-03 at 20:08:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 20:08:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: why was site 4 compared differently +2025-04-03 at 20:08:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Positions are extrapolated. + +The ll photographs have been normalized so that the earth is the same size in all frames. Frames 8590 and 8591 have been further enlarged. By viewing these two frames under a stereoscope, pronounced apparent relief is seen in the cloud patterns. The relief is so pronounced, in fact, that it cannot be attributed solely to height differences of clouds. It appears to result, in part, from the relative horizontal motion in the cloud fields; that is, clouds moving in the same direction as the spacecraft appear farther away than those moving in the direction opposite that of the spacecraftīŧŒ + +11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT +------ + +2025-04-03 at 20:08:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:08:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module post-tank anomaly maneuver +2025-04-03 at 20:08:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ + +2025-04-03 at 20:08:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apoll translunar free return trajectory parameters +2025-04-03 at 20:08:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +TABLE 4-II.- TRAJECTORY PARAMETERS +------ + +2025-04-03 at 20:08:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection aborted mission navigation tracking systems +2025-04-03 at 20:08:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 20:08:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: earth observation cloud movement +2025-04-03 at 20:08:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Positions are extrapolated. + +The ll photographs have been normalized so that the earth is the same size in all frames. Frames 8590 and 8591 have been further enlarged. By viewing these two frames under a stereoscope, pronounced apparent relief is seen in the cloud patterns. The relief is so pronounced, in fact, that it cannot be attributed solely to height differences of clouds. It appears to result, in part, from the relative horizontal motion in the cloud fields; that is, clouds moving in the same direction as the spacecraft appear farther away than those moving in the direction opposite that of the spacecraftīŧŒ + +11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT +------ +Result 2: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ + +2025-04-03 at 20:08:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:08:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module helium pressure regulation +2025-04-03 at 20:08:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:08:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 12 translunar non-free return velocity change +2025-04-03 at 20:08:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 20:08:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: transearth injection maneuver throttle profile descent propulsion system pitch roll attitude control +2025-04-03 at 20:08:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:08:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: cloud movement and third stage lunar impact +2025-04-03 at 20:08:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Positions are extrapolated. + +The ll photographs have been normalized so that the earth is the same size in all frames. Frames 8590 and 8591 have been further enlarged. By viewing these two frames under a stereoscope, pronounced apparent relief is seen in the cloud patterns. The relief is so pronounced, in fact, that it cannot be attributed solely to height differences of clouds. It appears to result, in part, from the relative horizontal motion in the cloud fields; that is, clouds moving in the same direction as the spacecraft appear farther away than those moving in the direction opposite that of the spacecraftīŧŒ + +11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:08:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:08:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 helium pressure rise rate investigation +2025-04-03 at 20:08:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:08:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: non-free return translunar profile comparisons +2025-04-03 at 20:08:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ + +2025-04-03 at 20:08:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: descent propulsion system free-return trajectory midcourse correction control systems +2025-04-03 at 20:08:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:08:54 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: S-IVB stage lunar impact site 4 +2025-04-03 at 20:08:54 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 20:08:54 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:08:54 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:08:54 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, False, False, False, False, False] +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.00 +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.00 +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_correctness:82 - Student lengths: [1851, 902, 1156, 554, 1965, 1857] +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [22, 22, 22, 22, 22, 22] +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_correctness:84 - Average student length: 1380.83 +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 22.00 +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_correctness:86 - Length ratio: 62.77 +2025-04-03 at 20:08:54 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.333, Valid formats: 2.0/6 +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 20:08:54 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.767 Âą 0.347 +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 7.33 Âą 3.94 +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [10, 4, 10, 0, 10, 10] +2025-04-03 at 20:08:54 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target. + +10.2 NETWORK +------ +Result 2: +The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel. + +The final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point. +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always +------ +Result 2: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency. + +e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived. + +The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented. +------ +Result 2: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:08:54 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nA major flight control function, in adaition to the monitoring of systems sta...', 'Result 1:\nA major flight control function, in adaition to the monitoring of systems sta...', 'Result 1:\nA major flight control function, in adaition to the monitoring of systems sta...', 'Result 1:\nPowering down of the command and service nodules and powering up of the lunar...', 'Result 1:\nd. The effectiveness of preflight crew training, especially in conjunction wi...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...'] +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Tables D-I and D-II are summaries of the data made available for systems performance analyses and anomaly investigations. Table D-l lists the data from the command and service modules, and table D-II, for the lunar module. For additional information regarding data availability, the status listing of all mission data in the Central Metric Data File, building l2, MSC, should be consulted. + +TABIE D-I.- COMMAND AND SERVICE MODULE DATA AVAILABILITY +------ +Result 2: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The command module arrived at the contractor's facility in Downey,' California, on April 27, 1970, after reaction control system deactivation and pyrotechnic safing in Hawaii. Postflight testing and inspection of the command module for evaluation of the inflight performance and investigation of the flight irregularities were conducted at the contractor's and vendors' facilities in accordance with approved Apollo Spacecraft Hardware Utilization Requests (ASHUR's). The tests performed as a result of inflight problems are described in table C-I and discussed in the appropriate systems performance sections of this report. Tests being conducted for other purposes in accordance with other ASHUR's and the basic contract are not included. + +TABLE C-I.- POSTFLIGHT TESTING SUMMARY +------ +Result 2: +A.2 LUNAR MODULE +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7. + +A complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. +------ +Result 2: +Temperature changes were noted in bays 3 and 4 of the service module in response to a high heat pulse or high pressure surge. Fuel cell 2 was turned off about 2 hours later because of the loss of pressure from cryogenic oxygen tank l. + +The cryogenic oxygen tank design will be changed to eliminate the mechanisms which could initiate burning within the tank and ultimately lead to a structural failure of the tank or its components. All electrical wires will be stainless-steel sheathed and the quantity probe will be made from stainless steel instead of aluminum. The fill-line plumbing internal to the tank will be improved, and a means of warning the crew of an inadvertent closure of either the fuel cell hydrogen or oxygen valves will be provided. A third cryogenic oxygen tank will be added to the service module for subsequent Apollo missions. The fuel cell oxygen +------ + +2025-04-03 at 20:08:54 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nTables D-I and D-II are summaries of the data made available for systems perf...', "Result 1:\nThe command module arrived at the contractor's facility in Downey,' Californi...", 'Result 1:\nBecause an inflight anomaly in the cryogenic oxygen supply required an abort ...', 'Result 1:\nBecause an inflight anomaly in the cryogenic oxygen supply required an abort ...'] +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Upon completion of photography, the two docked vehicles were maneuvered back to the service module separation attitude in preparation for the command module alignment. Star observation through the command module optics in this attitude was poor because of light reflecting from the lunar module, and the Conmander varied the pitch attitude by approximately 20 degrees in an attempt to improve star visibility. These attitude excursions, however, were not effective, and the spacecraft was returned to the original separation attitude for the command module alignment. + +8.9.5 Command Module Activation +------ +Result 2: +Following translunar injection, earth weather photography was conducted for approximately 6 hours. + +The first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget. +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds + +8.7 TRANSLUNAR FLIGHT + +8.7.1 Coast Phase Activities +------ +Result 2: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency , + + + +Figure 8-l.- Flight plan activities. + + + + + +Figure 8-l.- Continued + + + +(c) 69 to 122 hours. Figure 8-l.- Continued. + + + +(a) 122 to 143 hours. Figure 8-l.- Concluded. + +low-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value. + +8.6 TRANSPOSITION AND DOCKING +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +TABLE 4-II.- TRAJECTORY PARAMETERS +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II. + +TABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS +------ +Result 2: +data. Following this maneuver, a series of earth photographs were taken for later use in determining wind profiles in the upper atmosphere. The first midcourse correction inserted the spacecraft into a non-free-return trajectory. +------ + +2025-04-03 at 20:08:54 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nupdated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 ...', 'Result 1:\ndata. Following this maneuver, a series of earth photographs were taken for l...', 'Result 1:\nUpon completion of photography, the two docked vehicles were maneuvered back ...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nFollowing separation and translation, a manual pitch maneuver of 1.5 deg/sec ...', 'Result 1:\nNominal first-opportunity translunar injection procedures were used and are s...', 'Result 1:\nAs on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return t...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nAs on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return t...'] +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 20:08:54 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: [] +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +(section ll.3). +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +updated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 at 32:04:29 dupdated to -2.9 at 32:04:29 + +coefficient for the X-axis, which was still within specified limits; this coefficient being the most sensitive contributor to the gyrocompassing misalignment. Table 5.6-IlI is a set of error sources which reproduce the velocity errors observed during as cent. +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ +Result 2: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Table 5.6-I is a summary of gyro drift measurements deduced from inflight alignments. The null-bias drift coefficients for all three gyros were updated at 32 hours , based upon drift rates calculated from four platform alignments. The alignment prior to entry was performed by first conducting a coarse alignment to the lunar module platform and then using the automatic optics positioning capability to locate stars for a precise alignment. This technique was necessary because of the difficulty in recognizing constellations through the scanning telescope as a result of reflections from the lunar module and obscuration by vented particles. + +TABLE 5.6-I.\~ PLATFORM ALIGNMENT SUMMARY +------ +Result 2: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +To assure the alignment accuracy of the lunar module platform for the transearth injection maneuver, a check was made at 74 hours utilizing the sun for reference. The method involved a platform alignment program (P52, option 3), loading the sun vectors, and utilizing an automatic attitude maneuver. The null point was approximately one-half a sum diameter to the right of the sun's edge. A.two-diameter offset was allowable, So the platform was considered acceptable. +------ +Result 2: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees . + +Spacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12. + +(a) Trans lunar +------ +Result 2: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs. + +6.4.3 Alignment +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in "Auto." + +Primary guidance system performance was nomi nal $\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver. +------ +Result 2: +The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained. +------ + +2025-04-03 at 20:08:54 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nupdated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 ...', 'Result 1:\nupdated to -0.167 at 141:30:00 Pupdated to +0.6 at 32:0h:29 Cupdated to -1.2 ...', 'Result 1:\nTable 5.6-I is a summary of gyro drift measurements deduced from inflight ali...', 'Result 1:\nTable 5.6-II summarizes the inertial component preflight histories. Velocity ...', 'Result 1:\nTable 5.6-I is a summary of gyro drift measurements deduced from inflight ali...', 'Result 1:\nTable 5.6-I is a summary of gyro drift measurements deduced from inflight ali...', 'Result 1:\nTo assure the alignment accuracy of the lunar module platform for the transea...', 'Result 1:\nThe first transearth midcourse correction (table 4-III), was performed at 105...', 'Result 1:\nThe transearth injection maneuver was performed with the primary guidance sys...', 'Result 1:\nA descent propulsion system maneuver to reestablish a free-return trajectory ...'] +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary. + +5.1 SITRUCTURAL AND MECHANICAL SYSTEMS +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Although the standard format was followed during the deactivation and postrecovery inspection of the command module, it should be noted that extreme caution was taken during these operations to insure the integrity of the command module for postflight evaluation of the anomaly. After deactivation, the command module was secured and guarded. + +The following discrepancies were noted during the postrecovery inspection: +------ +Result 2: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by + + + +Figure ll.l-l.- Field meter location in the laumch site area. + + + +Figure ll.l-2.-- Field meter locations in the proximity of the launch complex. + +gravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6. +------ +Result 2: +The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12. +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +As a result of the electrical disturbances experienced during the Apollo l2 launch, the value of further research in this area was recognized and several experiments were performed prior to and during the Apollo l3 launch to study certain aspects of launch-phase electrical phenomena. The separate experiments consisted of measurements of the atmospheric electric field, low-frequency and very-low-frequency radio noise, the air/earth current density, and the electrical current flowing in the earth's surface, all of which result from perturbations generated by the launch vehicle and its exhaust plume. The analysis of the Apollo l2 lightning incident is reported in reference 3. + +11.1.1 Electric Field Measurements +------ +Result 2: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Illustrative data from the field instruments during launch are shown in figure ll.l-3. Very large perturbations of the normal electric field were recorded on meters at sites l, 2, and 3 located near the launch tower. First, there was a rapid increase in the positive direction, followed by a slower negative decrease. Data taken at site 4, however, did not indicate any significant variations in field intensity. Excellent records at several sensitivity levels were obtained at site 7. The field perturbation immediately following launch rose to a maximum of 1200 volts/ meter in about 25 seconds. The direction of field change then reversed, and a negative peak of some 300 volts/meter was reached in about ll5 seconds. Thereafter, the field gradually returned to the unperturbed value. +------ +Result 2: +As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges. +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ +Result 2: +Positions are extrapolated. + +The ll photographs have been normalized so that the earth is the same size in all frames. Frames 8590 and 8591 have been further enlarged. By viewing these two frames under a stereoscope, pronounced apparent relief is seen in the cloud patterns. The relief is so pronounced, in fact, that it cannot be attributed solely to height differences of clouds. It appears to result, in part, from the relative horizontal motion in the cloud fields; that is, clouds moving in the same direction as the spacecraft appear farther away than those moving in the direction opposite that of the spacecraftīŧŒ + +11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +Positions are extrapolated. + +The ll photographs have been normalized so that the earth is the same size in all frames. Frames 8590 and 8591 have been further enlarged. By viewing these two frames under a stereoscope, pronounced apparent relief is seen in the cloud patterns. The relief is so pronounced, in fact, that it cannot be attributed solely to height differences of clouds. It appears to result, in part, from the relative horizontal motion in the cloud fields; that is, clouds moving in the same direction as the spacecraft appear farther away than those moving in the direction opposite that of the spacecraftīŧŒ + +11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT +------ +Result 2: +gathered from pairs of geosynchronous satellites located l0 to 20 degrees apart in longitude. Calculations indicate, however, that stereoscopic determination of cloud heights from geosynchronous aititudes would be marginal, at best, because of the small disparity angles involved (ref.6). +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +Positions are extrapolated. + +The ll photographs have been normalized so that the earth is the same size in all frames. Frames 8590 and 8591 have been further enlarged. By viewing these two frames under a stereoscope, pronounced apparent relief is seen in the cloud patterns. The relief is so pronounced, in fact, that it cannot be attributed solely to height differences of clouds. It appears to result, in part, from the relative horizontal motion in the cloud fields; that is, clouds moving in the same direction as the spacecraft appear farther away than those moving in the direction opposite that of the spacecraftīŧŒ + +11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT +------ +Result 2: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ + +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0. +------ +Result 2: +Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11īŧŒ1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18 +------ + +2025-04-03 at 20:08:54 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: ['Result 1:\nAlthough the standard format was followed during the deactivation and postrec...', 'Result 1:\nAt site 6, the record was similar to that for site 7 with an initial positive...', 'Result 1:\nIllustrative data from the field instruments during launch are shown in figur...', 'Result 1:\nIllustrative data from the field instruments during launch are shown in figur...', 'Result 1:\nAs a result of the electrical disturbances experienced during the Apollo l2 l...', 'Result 1:\nIllustrative data from the field instruments during launch are shown in figur...', 'Result 1:\nMi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-...', 'Result 1:\nPositions are extrapolated.\n\nThe ll photographs have been normalized so that ...', 'Result 1:\nPositions are extrapolated.\n\nThe ll photographs have been normalized so that ...', 'Result 1:\nThe discarded S-IVB stage was targeted for a lunar impact of 3 degrees south ...'] +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 20:08:54 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.300, Perfect scores: 0/6 +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 3/6 +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 3.33 +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 20:08:54 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:08:54 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.614, Max reward: 0.845 +2025-04-03 at 20:08:56 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:08:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias update +2025-04-03 at 20:08:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:08:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "accelerometer bias update Z-axis value" +2025-04-03 at 20:08:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:08:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: what is the typical update frequency and documentation for Z-axis accelerometer bias +2025-04-03 at 20:08:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:08:59 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias update Z-axis acceleration +2025-04-03 at 20:08:59 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 20:08:59 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:09:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias update -1.42 +2025-04-03 at 20:09:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:09:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Z-axis accelerometer bias after update +2025-04-03 at 20:09:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:09:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: accelerometer bias update Z-axis sensor measured value +2025-04-03 at 20:09:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 20:09:01 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias after update +2025-04-03 at 20:09:01 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:09:01 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:09:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias update new value +2025-04-03 at 20:09:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:09:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Z-axis accelerometer bias values +2025-04-03 at 20:09:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:09:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: UPDATE Z-axis accelerometer bias 6.5 REACTION CONTROL +2025-04-03 at 20:09:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 20:09:03 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias update Z-axis acceleration +2025-04-03 at 20:09:03 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 20:09:03 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:09:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias update mission log +2025-04-03 at 20:09:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 20:09:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Z-axis accelerometer bias values 6.5 REACTION CONTROL +2025-04-03 at 20:09:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 20:09:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Latest recorded Z-axis accelerometer bias after update from 6.5 REACTION CONTROL +2025-04-03 at 20:09:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 20:09:05 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias after translunar injection +2025-04-03 at 20:09:05 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-03 at 20:09:05 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:09:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias update 6.4.5 +2025-04-03 at 20:09:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:09:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Z-axis accelerometer bias after instrumentation +2025-04-03 at 20:09:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:09:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Results of several entry monitor system bias tests including Z-axis acceleration bias during translation +2025-04-03 at 20:09:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ + +2025-04-03 at 20:09:07 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias after translunar injection +2025-04-03 at 20:09:07 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-03 at 20:09:08 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:09:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Adapter accelerometer bias after instrumentation +2025-04-03 at 20:09:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 20:09:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: attitude error bias after final midcourse correction +2025-04-03 at 20:09:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +initial maneuver to the firing attitude for the final midcourse correction was done manually using the earth as a reference in the same manner as the previous maneuver. This procedure presented no problems , even though the earth disk was considerably larger at this time. +------ + +2025-04-03 at 20:09:10 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: z-axis accelerometer bias after translunar injection 0.008 +2025-04-03 at 20:09:10 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-03 at 20:09:10 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:09:11 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: first and second midcourse correction attitude error bias +2025-04-03 at 20:09:11 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-03 at 20:09:11 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:09:11 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:09:11 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 1/6 answers correct +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, False, False, False, False] +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.17 +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.37 +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_correctness:82 - Student lengths: [280, 155, 1854, 2002, 1854, 269] +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [44, 44, 44, 44, 44, 44] +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_correctness:84 - Average student length: 1069.00 +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 44.00 +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_correctness:86 - Length ratio: 24.30 +2025-04-03 at 20:09:11 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.000, Valid formats: 0.0/6 +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 20:09:11 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.633 Âą 0.448 +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 4.00 Âą 2.89 +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [5, 0, 6, 7, 6, 0] +2025-04-03 at 20:09:11 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ +Result 2: +Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:09:11 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: ['Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nThe inertial measurement unit performed properly throughout the mission. A pr...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...'] +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-03 at 20:09:11 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: [] +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 20:09:11 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...'] +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +initial maneuver to the firing attitude for the final midcourse correction was done manually using the earth as a reference in the same manner as the previous maneuver. This procedure presented no problems , even though the earth disk was considerably larger at this time. +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +After completing the maneuver to the attitude required for the final midcourse correction, the crew reported that the attitude error needles were not nulled on the flight director attitude indicator. The sequence used to power up the platform and to enable the autopilot prevented certain computer memory cells from being properly initialized. Consequently, an attitude error bias was introduced between the stored values of attitude error and those displayed on the attitude error needles. When the digital autopilot is turned on, a computer routine checks the status of an "error counter enable" bit to see if initialization is required. If this bit is off, as it normally would be, initialization takes place and the error counter, certain memory cells, and the inertial coupling display umit digital-to-analog converters are all zeroed. If the computer check finds the error counter enabled, the assumption is made that initialization has already taken place and the calculated attitude error is +------ +Result 2: +The error counters for the coupling display units are used by the digital autopilot for attitude error displays, but are also used to drive the platform during a coarse alignment. A platform coarse alignment was performed at about l35 hours, and the error-counter-enable status bit was set. The digital autopilot was activated 2 hours later, but with the error cournters already enabled, no initialization took place and a bias was introduced into the attitude error loop. The attitude errors displayed to the crew at the completion of the attitude maneuver prior to the seventh midcourse correction reflected a bias in the pitch, roll, and yaw axes of plus l.3, plus 2l.2, and minus l2.0 degrees, respectively. + +Spacecraft dynamics were very small during the service module jettison and lunar module undocking sequence. Velocity changes imparted to the respective vehicles during each maneuver were as follows: + +6.4.2 Translation Maneuvers +------ + +2025-04-03 at 20:09:11 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nUncompens ated Error term error One-sigma specification Offset velocity, ft/s...', 'Result 1:\nAfter completing the maneuver to the attitude required for the final midcours...', 'Result 1:\nAfter completing the maneuver to the attitude required for the final midcours...'] +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table. + +Sample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axisīŧŒmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.¡ -5.38 2.37 4 -5.5 -4.0 + +6.4.5 Abort Guidance System Performance +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ +Result 2: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.¡ Y Z 96- 116 37 116 Lt- 116 Null bias driftīŧŒ mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5 + +Several entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table. + +5.7 REACTION CONTROL + +Time Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015 + +5.7.1 Servi ce Module +------ +Result 2: +Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 äēē -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample meeīŧŒ deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 äēē -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr + +6.5 REACTION CONTROL +------ + +2025-04-03 at 20:09:11 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nAccelerometer bias Sample me an, Standard deviation, Number of Bample8 Final ...', 'Result 1:\nUncompens ated Error term error One-sigma specification Offset velocity, ft/s...', 'Result 1:\nUncompens ated Error term error One-sigma specification Offset velocity, ft/s...', 'Result 1:\nUncompens ated Error term error One-sigma specification Offset velocity, ft/s...'] +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +2025-04-03 at 20:09:11 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 20:09:11 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.000, Perfect scores: 0/6 +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 0/6 +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.00 +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 0/6 +2025-04-03 at 20:09:11 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:09:11 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.267, Max reward: 0.745 +2025-04-03 at 20:09:13 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:09:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Curiosity rover fuel cell current +2025-04-03 at 20:09:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-03 at 20:09:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "National Aeronautics and Space Administration Shuttle fuel cell average current" +2025-04-03 at 20:09:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:09:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fuel cell average current in space mission +2025-04-03 at 20:09:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 20:09:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fuel cell efficiency apollo 11 +2025-04-03 at 20:09:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-03 at 20:09:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: power cell specifications curiosity rover +2025-04-03 at 20:09:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 20:09:17 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Mars 2020 Perseverance rover fuel cell power system nominal current" +2025-04-03 at 20:09:17 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 20:09:17 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:09:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fuel cell current ratings NASA +2025-04-03 at 20:09:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 20:09:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fuel cell power output and voltage +2025-04-03 at 20:09:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:09:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: average fuel cell current mars sample return mission +2025-04-03 at 20:09:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 20:09:20 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Spaceworks 8 propulsion fuel cell average current" +2025-04-03 at 20:09:20 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 20:09:20 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:09:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fuel cell performance NASA Space Shuttle +2025-04-03 at 20:09:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal. + +During the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l. +------ + +2025-04-03 at 20:09:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: average hydrogen (fuel cell) current power supply mission +2025-04-03 at 20:09:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 20:09:23 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Helion fuel cell average current" +2025-04-03 at 20:09:23 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:09:23 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:09:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: ground fuel cell current supply unscheduled shutdowns +2025-04-03 at 20:09:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:09:25 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Mars 2020 rover fuel cell 1 average current" +2025-04-03 at 20:09:25 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 20:09:25 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:09:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: fuel cell shutdown current NASA Space Shuttle +2025-04-03 at 20:09:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal. + +During the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l. +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:09:27 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Mars 2020 rover fuel cell electrical discharge" +2025-04-03 at 20:09:27 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-03 at 20:09:27 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:09:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Space Shuttle average fuel cell operating current +2025-04-03 at 20:09:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:09:29 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: "Mars 2020 rover fuel cell normal operation current" +2025-04-03 at 20:09:29 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-03 at 20:09:29 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:09:31 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: average current Space Shuttle fuel cells +2025-04-03 at 20:09:31 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:09:31 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:09:32 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:09:32 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:09:32 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 5/6 answers correct +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [True, False, True, True, False, True] +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.67 +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_correctness:82 - Student lengths: [288, 109, 453, 218, 310, 199] +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [10, 10, 10, 10, 10, 10] +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_correctness:84 - Average student length: 262.83 +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 10.00 +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_correctness:86 - Length ratio: 26.28 +2025-04-03 at 20:09:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.500, Valid formats: 3.0/6 +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_format:228 - Responses ending properly: 4/6 +2025-04-03 at 20:09:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.546 Âą 0.207 +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 4.00 Âą 3.56 +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 1/6 +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [1, 11, 2, 1, 3, 6] +2025-04-03 at 20:09:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal. + +During the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l. +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +Efforts to install the tunnel hatch were terminated when the Commander observed venting of material from the service module area. He then reported the oxygen tank 2 pressure was zero and oxygen tank l pressure was decreasing. This information pinpointed the problem source to within the command and service modules. + +At ground request, fuel cells l and 3 regulator pressures were read from the systems test meter, confirming the loss of these fuel cells. AC bus 2 was tied. to inverter 1, and the emergency power-down procedure was initiated to reduce the current flow to l0 amperes. At ground request, fuel cell l and, shortly thereafter, fuel cell 3 were shutdown in an attempt to stop the decrease in oxygen tank l pressure. +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal. + +During the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l. +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ +Result 2: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries. + +5.2.2 Fuel Cells +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The electrical power system performed all required functions. At lunar module undocking, the descent batteries had delivered 1434.7 amperehours from a nominal total capacity of l6o0 ampere-hours, and the ascent batteries had delivered 200 ampere-hours from a nominal total of 592 ampere-hours. The lunar module initial powered-down configuration required an average electrical energy consumption of 900 watts at 30 amperes. After the second descent propulsion firing, the lunar module was further powered down to about a 360-watt (l2-armpere) level; as discussed in section 7.2. A false battery 2 malfunction and master alarm occurred at 99:54:00 and continued intermittently during the perioas that the battery was on (discussed in section l4.2.3). A review of the data indicates that a current surge of greater than 100 amperes occurred at 97:13:56 concurrent with a crew report of a thumping noise and snowflakes seen through the lunar module window. This occurrence is discussed in section +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +The prelaunch timeline was satisfactory, and adequate time was allotted for suiting and associated activities to meet the ingress time. The final count was smooth and communications with the Test Conductor and the Mission Control Center were adequate. After the fuel cell selector knob was rotated and had been in the new position for a short time, the fuel cell flow indicators would alternately rise several scale marks and then return to normal momentarily before cycling again. Since this effect was observed for all three fuel cells, the possibility of a sensor anomaly was dismissed. With the crew fully strapped down, some difficulty was encountered in removing the helmet protective covers just prior to egress of the closeout personnel. + +8.3 LAUN CH +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Following recovery of the data, the vehicle had experienced a translation change of about 0.4 ft/sec, primarily in a plane normal to bay 4. The oxygen tank 2 pressure indication was at the lower limit of the readout. The oxygen tank l heaters were on, and the tank l pressure was decaying rapidly. A main-bus-B undervoltage alarm and a computer restart also occurred at this time. + +Fuel cells l and 3 operated for about 2-l/2 minutes after the reactant valves closed. During this period, these fuel cells consumed the oxygen trapped in the plumbing, thereby reducing the pressure below minimum requirements and causing total loss of fuel cell current and voltage output from these two fuel cells. Because of the loss of performance by two of the three fuel cells and the subsequent load switching by the crew, numerous associated master alarms occurred as expected. +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to + +operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking. + + + + + +Figure 7.2-2.- Lunar module total battery capacity during flight. + + + +Apollo 13 flight crew +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts. + +5.3 CRYOGENIC STORAGE +------ +Result 2: +Approximately 2 seconds later, the Command Module Pilot reported a master alarm and a main-bus-B undervoltage light. Voltage readouts from main bus B, fuel cell 3 current, and reactant flows were normal, and it was concluded a transient had occurred. The Command Module Pilot then initiated efforts to install the tunnel hatch. + +The Lunar Module Pilot proceeded to the right seat and found the ac-bus-2 and ac-bus-2-overload warning lights on, with main bus B voltage, fuel-cell-3 current, and fuel-ceil-3 reactant flow indications offScale low. Inverter 2 was then removed from main bus B. + +On switching ac electrical loads to ac bus l, the main bus A undervoltage light illuminated, with a corresponding reading of 25.5 volts. A check of the fuel cells revealed fuel cell l reactant flow to be zero. At all times, fuel cells l and 2 were tied to main bus A and fuel cell 3 to main bus B, with the. proper grey flags displayed. +------ + +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 6.0 +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 1.000 +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 20:09:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.433, Perfect scores: 0/6 +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 4/6 +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 1.00 +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 3/6 +2025-04-03 at 20:09:32 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:09:32 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.314, Max reward: 0.642 +2025-04-03 at 20:09:34 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts +2025-04-03 at 20:09:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: How long are electronic devices typically cold soaked to maintain efficiency +2025-04-03 at 20:09:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-03 at 20:09:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature tolerance cold soak duration 7 hours +2025-04-03 at 20:09:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 20:09:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature for cold soak in semiconductor process 7 hours +2025-04-03 at 20:09:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-03 at 20:09:36 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: query. + + temperature testing lithium-ion batteries 7-hour cold soak +2025-04-03 at 20:09:36 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +Evidence indicates that battery 2 may have experienced an electrical fault of some type. The most probable condition is electrolyte leaking from one or more cells and bridging the high-voltage or low-voltage terminal to the battery case (fig. 14-17). This bridging results in water electrolysis and subsequent ignition of the hydrogen and oxygen so generated. The accompanying "explosion" would then blow off or rupture the seal of the battery lid and cause both a thump and venting of the free liquids in the battery case, resulting in "snowflakes." + +Postflight tests have shown the following: +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 20:09:36 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:09:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module cold soak temperature +2025-04-03 at 20:09:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 20:09:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: temperature cold soak Apollo 11 low pressure tolerance +2025-04-03 at 20:09:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:09:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 cold soak temperature +2025-04-03 at 20:09:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:09:38 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 crew temperature of liquid water tanks +2025-04-03 at 20:09:38 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 20:09:38 | INFO | src.agent:run_agent_generations:70 - Generating responses for 4 prompts +2025-04-03 at 20:09:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo lunar module cold soak temperature after power-down +2025-04-03 at 20:09:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 20:09:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: abnormally high pressure rise rate during Apollo 13 top-off procedure +2025-04-03 at 20:09:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:09:41 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 cold soak temperature range lunar module +2025-04-03 at 20:09:41 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 20:09:41 | INFO | src.agent:run_agent_generations:70 - Generating responses for 3 prompts +2025-04-03 at 20:09:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: lunar module helium pressure rise rate Apollo 13 cold soak adjustment +2025-04-03 at 20:09:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:09:43 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 11 cold soak temperature estimated range +2025-04-03 at 20:09:43 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:09:43 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:09:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 lunar module helium pressure rise rate management during cold soak +2025-04-03 at 20:09:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:09:45 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 cold soak helium temperature +2025-04-03 at 20:09:45 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:09:45 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:09:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 operating temperature +2025-04-03 at 20:09:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 20:09:47 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 preflight helium tank test temperature +2025-04-03 at 20:09:47 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:09:47 | INFO | src.agent:run_agent_generations:70 - Generating responses for 2 prompts +2025-04-03 at 20:09:48 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 helium tank cold soak temperature estimate +2025-04-03 at 20:09:48 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:09:48 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:09:50 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 cold soak temperature helium 900 psia +2025-04-03 at 20:09:50 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:09:50 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:09:51 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 cold soak temperature replicate test 900 psia +2025-04-03 at 20:09:51 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:09:51 | INFO | src.agent:run_agent_generations:70 - Generating responses for 1 prompts +2025-04-03 at 20:09:52 | INFO | src.agent:run_tool_calls:115 - 🔍 Search Query: Apollo 13 cold soak helium pressure 900 psia temperature +2025-04-03 at 20:09:52 | INFO | src.agent:run_tool_calls:118 - ℹī¸ Information: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:09:52 | INFO | src.evaluation:check_student_answers:57 - Checking 6 student answers +2025-04-03 at 20:09:52 | INFO | src.evaluation:check_student_answers:83 - Generating verification responses +2025-04-03 at 20:09:52 | INFO | src.evaluation:check_student_answers:103 - Verification complete. 3/6 answers correct +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_correctness:75 - Correctness metrics: [False, True, True, False, False, False] +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_correctness:76 - Average correctness: 0.33 +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_correctness:77 - Standard deviation: 0.47 +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_correctness:82 - Student lengths: [572, 312, 134, 1851, 255, 487] +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_correctness:83 - Teacher lengths: [5, 5, 5, 5, 5, 5] +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_correctness:84 - Average student length: 601.83 +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_correctness:85 - Average teacher length: 5.00 +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_correctness:86 - Length ratio: 120.37 +2025-04-03 at 20:09:52 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_correctness.jsonl +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_format:227 - Format reward metrics - Mean: 0.167, Valid formats: 1.0/6 +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_format:228 - Responses ending properly: 3/6 +2025-04-03 at 20:09:52 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_format.jsonl +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_retry:312 - Retry behavior rewards: 0.508 Âą 0.393 +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_retry:313 - Search tags per completion: 3.50 Âą 3.55 +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_retry:314 - Violations (>1 search per message): 0/6 +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_retry:315 - Search counts distribution: [0, 3, 6, 10, 2, 0] +2025-04-03 at 20:09:52 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_retry.jsonl +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-03 at 20:09:52 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 0: +Search results: [] +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 20:09:52 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 1: +Search results: ['Result 1:\nDuring the period when the command module was powered down, the cabin tempera...', 'Result 1:\nAll mechanical systems functioned properly. One mechanical anomaly, however, ...', 'Result 1:\nAll mechanical systems functioned properly. One mechanical anomaly, however, ...'] +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ +Result 2: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ + +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ + +2025-04-03 at 20:09:52 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 2: +Search results: ['Result 1:\nDuring the period when the command module was powered down, the cabin tempera...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...'] +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +the cold soak. These vapors would slowly condense on the cold( $...0^{\circ}\mathrm{~\textsf~{~R~}~})$ inner wall,resulting in the pressure rise rate droping to the nominal level, as was observed. The rise rate would remain normal until the helium temperature increased above the vaporization temperature of the contaminant. +------ +Result 2: +Earlier, at approximately 73 hours, the command module windows had become nearly opaque with water droplets. This moisture contamination continued to increase, and at approximately llo hours a thin wafer film appeared on the interior command module structure itself, as well as on the lunar module windows. Despite this condensation because of the reduced cabin temperature, at no time did the humidity reach levels which were uncomfortable to the crew. The moisture on the lunar module windows disappeared shortly after power-up at approximately l35 hours. The condensation generally disappeared after parachute deployment, although the structure remained cold even after landing. +------ + +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 3: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +All mechanical systems functioned properly. One mechanical anomaly, however, was a gas leak from one of two breech assemblies in the apex cover jettison system, and this problem is discussed in section 14.l.6. In addition, docking timnel insulation, which normally remains with the lunar module after separation, was noted from photographs to have cracked and expanded radially. Since the cracking is believed to occur during pyrotechnic firing and has been seen in past flights, it is not a problem. + +Structural temperatures remained within acceptable limits throughout the mission. However, because of the long cold-soak period following powering down, the comnand module structure exhibited significantly lower temperatures than has been observed in previous flights. + +5.2 ELECTRICAL POWER + +5.2.1 Batteries +------ + +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 4: Result 1: +After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\circ}$ Or $60^{\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\mathsf{c m}/\mathsf{s e c}^{2}$ to the new value of minus $1.66~\mathsf{c m}/\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments . +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 5: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 6: Result 1: +helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively. + +The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank. +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 7: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 8: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ + +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 9: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 10: Result 1: +During the initial cold-soak period following loading of supercritical helium during the Apollo l3 countdown demonstration test, the helium exhibited a pressure rise rate approximately three times greater than expected. A preflight test was devised to determine the pressure-rise rate that would exist at the time of descent engine firing for lunar descent. The predicted tank conditions at that time would be approximately 900 psia pressure and $48$ pounds of helium. Normal procedures were not used to reach 900 psia because 1o0 hours would have been required and the launch schedule would have been impacted; therefore, the pressure was raised to 900 psia by flowing warm helium through the tank heat exchanger. The subsequent pressure rise rate was abnormally high at l4.9 psi/hour. The abnormality of this rate was confirmed by repeating the test on two other +------ +Result 2: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ + +2025-04-03 at 20:09:52 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 3: +Search results: ['Result 1:\nthe cold soak. These vapors would slowly condense on the cold( $...0^{\\circ}\\...', 'Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...', 'Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...', 'Result 1:\nAfter the oxygen tank incident, the platform was used as a reference to which...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', "Result 1:\nhelium tanks, one at the manufacturer's plant and the other at the Manned Spa...", 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...', 'Result 1:\nDuring the initial cold-soak period following loading of supercritical helium...'] +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 1: Result 1: +Evidence indicates that battery 2 may have experienced an electrical fault of some type. The most probable condition is electrolyte leaking from one or more cells and bridging the high-voltage or low-voltage terminal to the battery case (fig. 14-17). This bridging results in water electrolysis and subsequent ignition of the hydrogen and oxygen so generated. The accompanying "explosion" would then blow off or rupture the seal of the battery lid and cause both a thump and venting of the free liquids in the battery case, resulting in "snowflakes." + +Postflight tests have shown the following: +------ +Result 2: +During the period when the command module was powered down, the cabin temperature slowly decreased to approximately 43° F and considerable amounts of moisture condensed on the spacecraft windows and the command module structure. Thermal control, after powering up at 14o hours, was Satisfactory, although the cabin temperature remained very cold during entry. The command module potable water served as the main drinking supply for the crew during the mission, and approximately l4 pounds were withdrawn after powering down, using the 8-ounce plastic bags. The crew reported at approximately l20 hours they were unable to withdraw water from the potable tank and assumed it was empty. Approximately 6 hours after landing, the recovery crew was also unable to obtain a water sample from either the potable or waste water tanks. The recovery personnel stated the structure near the tank and lines was very cold to touch, and an analysis of temperatures during the flight in this vicinity show that +------ + +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:368 - 🔍 Searched Chunk 2: Result 1: +The operational support provided by the flight control team was satisfactory and timely in safely returning the Apollo 13 crew. Only the inflight problems which influenced flight control operation and their resultant effects on the flight plan are discussed. + +Prior to laurch, the supercritical.helium pressure in the lunar module descent propulsion system increased at an abnormally high rate. After cold soak ard venting, the rise rate was considered acceptable for launch. At 56 hours during the first entry into the lunar module, the rise rate and pressure were reported to be satisfactory; therefore, a special venting procedure was not required. + +A master caution and warning alarm at 38 hours indicated the hydrogen tank pressures were low. As a result, it was planned to use the cryogenic tank fars more often than scheduled to provide a more even distribution of fluid and to stabilize heat and pressure rise rates. +------ +Result 2: +6.7 ENVIRONMENTAL CONTROL + +Environmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5. + +An indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr. +------ + +2025-04-03 at 20:09:52 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 4: +Search results: ['Result 1:\nEvidence indicates that battery 2 may have experienced an electrical fault of...', 'Result 1:\nThe operational support provided by the flight control team was satisfactory ...'] +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:366 - 📝 Ground Truth Chunk: An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive. + +A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests. + +Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated. +2025-04-03 at 20:09:52 | WARNING | src.rewards:reward_em_chunk:374 - Failed to find correct chunk for prompt 5: +Search results: [] +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:384 - Chunk Query Rewards Summary: +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:385 - Total prompts: 6 +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:386 - Correct matches: 0.0 +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:387 - Average reward: 0.000 +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_em_chunk:388 - Reward std: 0.000 +2025-04-03 at 20:09:52 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_em_chunk.jsonl +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_search_strategy:612 - Search strategy metrics - Mean: 0.167, Perfect scores: 1/6 +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_search_strategy:615 - Initial searches: 1/6 +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_search_strategy:616 - Average info processing steps: 0.50 +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_search_strategy:617 - Final synthesis rate: 1/6 +2025-04-03 at 20:09:52 | INFO | src.rewards:log_chat_state:837 - 💾 Appended 6 chat states to /workspace/DeepSearch/logs/chat_states/chat_states_search_strategy.jsonl +2025-04-03 at 20:09:52 | INFO | src.rewards:reward_search_diversity:793 - Search diversity metrics - Mean reward: 0.394, Max reward: 0.816 +2025-04-03 at 20:09:54 | INFO | src.agent:run_agent_generations:70 - Generating responses for 6 prompts