[
  {
    "chunk_id": 2,
    "question": "On what date did the mission PA-1 experience a first pad abort?",
    "answer": "Nov.7, 1963",
    "difficulty": "easy",
    "chunk_content": "Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966"
  },
  {
    "chunk_id": 2,
    "question": "Where did the mission AS-102 take place?",
    "answer": "Cape Kennedy, Fla.",
    "difficulty": "medium",
    "chunk_content": "Mission Spacecraft Description Laumch date Launch site PA-1 BP-6 First pad abort Nov.7, 1963 White Sands Missile Range; A-001 BP-12 Transonic abort May 13, 1964 N.Mex. White Sands Missile Range, AS-101 BP-13 Nominal launch and exit environment May 28, 1964 N. Mex. Cape Kennedy. Fla. AS-102 BP-15 Nominal launch and exit environment Sept.18,1964 Cape Kennedy, Fla. A-002 BP-23 Maximum dynamic pressure abort Dec.8, 1964 White Sands Missile Range, AS-103 BP-16 Micrometeoroid experiment Feb. 16, 1965 N.Mex. Cape Kennedy, Fla. A-003 BP-22 Low-altitude abort (planned high- May 19, 1965 White Sands Missile Range, AS-104 BP-26 altitude abort) Micrometeoroid experiment and service module May 25, 1965 N.Mex, Cape Kennedy, Fla. PA-2 BP-23A RCS launch environment Second pad abort June 29,1965 White Sands Missile Range, AS-105 BP-9A Micrometeoroid experiment and service module July 30, 1965 N. Mex. Cape Kennedy, Fla. A-004 SC-002 RCS launch environment Power-on tumbling boundary abort Jan.20,1966"
  },
  {
    "chunk_id": 3,
    "question": "Where was the AS-201 SC-009 Supercircular entry located on February 26, 1966?",
    "answer": "N. Mex.",
    "difficulty": "easy",
    "chunk_content": "White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla."
  },
  {
    "chunk_id": 3,
    "question": "What type of entry did the AS-202 SC-011 experience on August 25, 1966?",
    "answer": "Supercircular entry with high heat load",
    "difficulty": "medium",
    "chunk_content": "White Sands Missile Range, AS-201 SC-009 Supercircular entry with high heat rate Feb. 26,1966 N. Mex. Cape Kennedy, Fla. AS-202 SC-011 Supercircular entry with high heat load Aug.25,1966 Cape Kennedy, Fla."
  },
  {
    "chunk_id": 4,
    "question": "What is the title of the report?",
    "answer": "APOLLO 13 MISSION REPORT",
    "difficulty": "easy",
    "chunk_content": "MSC-02680\n\nCHANGE SHEET\n\nFOR\n\nNASA-MSC INTERNAL REPORT\n\nAPOLLO 13 MISSION REPORT\n\nChange 1\n\n\n\nMay 1970\n\nJames A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program\n\nAfter the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover \"Change l inserted.\"\n\nIn addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated.\n\nNOTE: A black bar in the margin of affected pages indicates the information that was changed or added.\n\n7.1.6 Batteries"
  },
  {
    "chunk_id": 4,
    "question": "Who is the Manager, Apollo Spacecraft Program?",
    "answer": "James A. MeDivitt Colonel, USAF",
    "difficulty": "easy",
    "chunk_content": "MSC-02680\n\nCHANGE SHEET\n\nFOR\n\nNASA-MSC INTERNAL REPORT\n\nAPOLLO 13 MISSION REPORT\n\nChange 1\n\n\n\nMay 1970\n\nJames A. MeDivitt Colonel, USAF Manager, Apollo Spacecraft Program\n\nAfter the attached enclosures (pages 7-3, 7-4, 7-7, 7-8, 11-3 through 11-6, E-3, E-4, and back cover), which are replacement pages, have been inserted, insert this CHANGE SHEET between the cover and title page and write on the cover \"Change l inserted.\"\n\nIn addition to the attached changes, please complete the attached Mission Report Questionaire and return as indicated.\n\nNOTE: A black bar in the margin of affected pages indicates the information that was changed or added.\n\n7.1.6 Batteries"
  },
  {
    "chunk_id": 5,
    "question": "How many ampere-hours of energy remained in the batteries at landing?",
    "answer": "29",
    "difficulty": "easy",
    "chunk_content": "The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained.\n\n\n\nFigure 7.l-l.- Entry battery energy.\n\n7.2 LUNAR MODULE\n\nFollowing lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of\n\n7.1.3 Cryogenic Fluids\n\nCryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls\u3002"
  },
  {
    "chunk_id": 5,
    "question": "How long did it take for oxygen tank 1 to be depleted after the incident?",
    "answer": "2 hours",
    "difficulty": "medium",
    "chunk_content": "The command module was completely powered down at 58 hours 40 minutes , at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained.\n\n\n\nFigure 7.l-l.- Entry battery energy.\n\n7.2 LUNAR MODULE\n\nFollowing lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of\n\n7.1.3 Cryogenic Fluids\n\nCryogenic oxygen and hydrogen usages were nominal until the time of the incident. The pressure decay in oxygen tank 2 was essentially instantaneous , while oxygen tank l was not depleted until approximately 2 hours following the incident. Usages listed in the following table are based on an analysis of the electrical power produced by the fuel ce1ls\u3002"
  },
  {
    "chunk_id": 7,
    "question": "What was the estimated total energy transferred to the command module?",
    "answer": "approximately 129 ampere hours",
    "difficulty": "easy",
    "chunk_content": "operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking.\n\n\n\nFigure 7.2-l.- Lunar module water usage.\n\n\n\nFigure 7.2-2.- Lunar module total battery capacity during flight.\n\n\n\nFigure ll.l-2.- Field meter locations in the proximity of the launch complex.\n\ngravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6."
  },
  {
    "chunk_id": 7,
    "question": "What was the amount of ampere hours remaining in the lunar module batteries at the time of undocking?",
    "answer": "410",
    "difficulty": "easy",
    "chunk_content": "operate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking.\n\n\n\nFigure 7.2-l.- Lunar module water usage.\n\n\n\nFigure 7.2-2.- Lunar module total battery capacity during flight.\n\n\n\nFigure ll.l-2.- Field meter locations in the proximity of the launch complex.\n\ngravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6."
  },
  {
    "chunk_id": 8,
    "question": "What was found on the moving portion of the records from sites 8 and 9?",
    "answer": "anything",
    "difficulty": "easy",
    "chunk_content": "Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off.\n\n\n\nFigure ll.l-3.- Electrical discharge data for the Apollo l3 launch.\n\n\n\nFigure ll.l-3.- Concluded\n\nNo significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations."
  },
  {
    "chunk_id": 8,
    "question": "At which stations were no significant perturbation in the electric field produced by the launch cloud?",
    "answer": "4 or 5",
    "difficulty": "medium",
    "chunk_content": "Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortunately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off.\n\n\n\nFigure ll.l-3.- Electrical discharge data for the Apollo l3 launch.\n\n\n\nFigure ll.l-3.- Concluded\n\nNo significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations."
  },
  {
    "chunk_id": 9,
    "question": "What type of signals were recorded at the launch site during the afternoon of launch day?",
    "answer": "Sporadic signals",
    "difficulty": "easy",
    "chunk_content": "The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12.\n\nThe above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood."
  },
  {
    "chunk_id": 9,
    "question": "What was the weather situation described as before the launch of the Apollo 13 vehicle?",
    "answer": "Marginal",
    "difficulty": "medium",
    "chunk_content": "The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12.\n\nThe above field meter records indicate the launch of the Apollo 13 vehicle produced a significant separation of electrical charge which could possibly increase the hazard in an otherwise marginal weather situation. At the present time the location and amount of the charge on the vehicle or exhaust clouds or a combination thereof are not well understood."
  },
  {
    "chunk_id": 10,
    "question": "What is the equilibrium potential that a conventional jet aircraft can approach?",
    "answer": "a million volts",
    "difficulty": "easy",
    "chunk_content": "It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more.\n\nSupplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation"
  },
  {
    "chunk_id": 10,
    "question": "What is the publication status of the document titled \"Preparati on 4 Ascent Propulsion System Final Flight Evaluation\"?",
    "answer": "Final Flight Evaluation",
    "difficulty": "easy",
    "chunk_content": "It is known that the electrostatic potentials develop on jet aircraft. These are caused by an engine charging current, which is balanced by a corona current loss from the aircraft. For.a conventional jet aircraft, the equilibrium potential can approach a million volts. For the Saturn V launch vehicle, the charging current may be larger than that of a jet aircraft, and therefore, the equilibrium potential for the Saturn vehicle might be on the order of a million volts or more.\n\nSupplement number Title Publication date/status Apollo 12 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight Evaluation September 1970 2 September 1970 3 Preparati on 4 Ascent Propulsion System Final Flight Evaluation 5 Descent Propulsion System Final Flight Preparation"
  },
  {
    "chunk_id": 11,
    "question": "What was the publication date of the report titled \"Trajectory Reconstruction and Analysis\"?",
    "answer": "March 1970",
    "difficulty": "easy",
    "chunk_content": "Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5"
  },
  {
    "chunk_id": 11,
    "question": "What was the status of the report titled \"Analysis of Apollo 10 Photography and Visual In\"?",
    "answer": "Cancelled",
    "difficulty": "easy",
    "chunk_content": "Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis March 1970 2 Guidance, Navigation, and Control System Performance Analysis December 1969 3 Performance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluation Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluati on Descent Propulsion System Final Flight Evaluation January 1970 8 9 Cancelled Analysis of Apollo l0 Photography and Visual In publication 10 Observations Entry Postflight Analysis December 1969 11 Communications System Performance December 1969 Apollo 1l 1 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module May 1970 September 1970 Reaction Control System Review Service Propulsion System Final Flight Evaluation Review 5"
  },
  {
    "chunk_id": 12,
    "question": "What was the subject of the document titled \"Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation\"?",
    "answer": "Ascent Propulsion System",
    "difficulty": "easy",
    "chunk_content": "Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970"
  },
  {
    "chunk_id": 12,
    "question": "In which month was the Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis conducted?",
    "answer": "December",
    "difficulty": "easy",
    "chunk_content": "Performance of Lunar Module Reaction Control System Review 6 Ascent Propulsion System Final Flight Evaluation September 1970 7 Descent Propulsion System Final Flight Evaluati on September 1970 8 Cancelled 9 10 11 Apollo ll Preliminary Science Report Communications System Performance Entry Postflight Analysis December 1969 January 1970"
  },
  {
    "chunk_id": 13,
    "question": "What year was the document approved by the Mission Evaluation Team?",
    "answer": "1970",
    "difficulty": "easy",
    "chunk_content": "PREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970\n\n\n\nTABLE OF CONTENTS\n\nSection Page"
  },
  {
    "chunk_id": 13,
    "question": "Where is the National Aeronautics and Space Administration Manned Spacecraft Center located?",
    "answer": "Houston, Texas",
    "difficulty": "easy",
    "chunk_content": "PREPARED BY\n\nMission Evaluation Team\n\nAPPROVED BY\n\n\n\nNATIONAL AERONAUTICS AND SPACE ADMINISTRATION MANNEDSPACECRAFTCENTER HOUSTON, TEXAS September 1970\n\n\n\nTABLE OF CONTENTS\n\nSection Page"
  },
  {
    "chunk_id": 14,
    "question": "What is the section that discusses the performance of the command and service module?",
    "answer": "5.0",
    "difficulty": "easy",
    "chunk_content": "1.0 SUMMARY 1-1 2.0 INTRODUCTION\u00b7. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... \u00b7\u00b7\u00b7 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .\u00b7 5-1 5.2 ELECTRICAL POWER \u00b7\u00b7\u00b7\u00b7 5-2 5.3 CRYOGENIC STORAGE.\u00b7\u00b7\u00b7 5-3 5.4 COMMUNICATIONS EQUIPMENT \u00b7 5-4 5.5 INSTRUMENTATION.\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .\u00b7 5-5 5.7 REACTION CONTROL.\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7 5-11 5.8 ENVIRONMENTAL CONTROL .\u00b7. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL \u00b7\u00b7\u00b7 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .\u00b7 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION \u00b7\u00b7\u00b7 6-8 6.7 ENVIRONMENTAL CONTROL.\u00b7\u00b7\u00b7 6-9 7.0 MISSION CONSUMABLES \u00b7\u00b7\u00b7\u00b7\u00b7. \u00b7\u00b7\u3001\u00b7 7-1 7.1 COMMAND AND SERVICE MODULES .\u00b7\u00b7\u00b7\u00b7 7-1 7.2 LUNAR MODULE \u00b7\u00b7\u00b7\u00b7\u00b7 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION \u2019\u00b7 8-2 8.6 TRANSPOSITION AND DOCKING .\u00b7.. 8-7"
  },
  {
    "chunk_id": 14,
    "question": "What is the section that describes the training of the pilots?",
    "answer": "8.1",
    "difficulty": "easy",
    "chunk_content": "1.0 SUMMARY 1-1 2.0 INTRODUCTION\u00b7. 2-1 3.0 MISSION DESCRIPTION 3-1 4.0 TRAJECTORY...... \u00b7\u00b7\u00b7 4-1 5.0 COMMAND AND SERVICE MODULE PERFORMANCE . . 5-1 5.1 STRUCTURAL AND MECHANICAL SYSTEMS .\u00b7 5-1 5.2 ELECTRICAL POWER \u00b7\u00b7\u00b7\u00b7 5-2 5.3 CRYOGENIC STORAGE.\u00b7\u00b7\u00b7 5-3 5.4 COMMUNICATIONS EQUIPMENT \u00b7 5-4 5.5 INSTRUMENTATION.\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7 5-4 5.6 GUIDANCE, NAVIGATION, AND CONTROL . .\u00b7 5-5 5.7 REACTION CONTROL.\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7 5-11 5.8 ENVIRONMENTAL CONTROL .\u00b7. 5-12 6.0 LUNAR MODULE PERFORMANCE 6-1 6.1 STRUCTURAL \u00b7\u00b7\u00b7 6-1 6.2 ELECTRICAL POWER 6-1 6.3 COMMUNICATIONS EQUIPMENT 6-2 6.4 GUIDANCE, NAVIGATION, AND CONTROL .\u00b7 6-2 6.5 REACTION CONTROL ... 6-8 6.6 DESCENT PROPULSION \u00b7\u00b7\u00b7 6-8 6.7 ENVIRONMENTAL CONTROL.\u00b7\u00b7\u00b7 6-9 7.0 MISSION CONSUMABLES \u00b7\u00b7\u00b7\u00b7\u00b7. \u00b7\u00b7\u3001\u00b7 7-1 7.1 COMMAND AND SERVICE MODULES .\u00b7\u00b7\u00b7\u00b7 7-1 7.2 LUNAR MODULE \u00b7\u00b7\u00b7\u00b7\u00b7 7-4 8.0 PILOTS' REPORT . . . 8-1. 8.1 TRAIN ING 8-1 8.2 PRELAUNCH PREPARATION .. 8-1 8.3 LAUN CH 8-2 8.4 EARTH ORBIT.. 8-2 Section Page 8.5 TRANSLUNAR INJECTION \u2019\u00b7 8-2 8.6 TRANSPOSITION AND DOCKING .\u00b7.. 8-7"
  },
  {
    "chunk_id": 15,
    "question": "What is the title of the section that discusses the physical examinations of the astronauts?",
    "answer": "9.3 PHYSICAL EXAMINATIONS",
    "difficulty": "easy",
    "chunk_content": "8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST \u00b7\u00b7\u00b7\u00b7\u00b7 8-11 8.10 ENTRY AND LANDING.\u00b7\u00b7. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY \u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7 \u00b7 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL \u00b7\u00b7\u00b7\u00b7 10-1 10.2 NETWORK\uff0e\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7 10-2 10.3 RECOVERY OPERATIONS...\u00b7.\u00b7. \u00b7\u00b7\u00b7 10-2 11.0 EXPERIMENTS\u00b7\u00b7\u00b7\u00b7\u00b7\uff0e\u00b7\u00b7\u00b7.\u00b7\u00b7\u00b7\u00b7. \u00b7\u00b7 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES\uff0e\u00b7\uff0e\u00b7\u00b7\uff0e\u00b7\uff0e\u00b7\u00b7\uff0e\u00b7\uff0e\u00b7\u00b7\uff0e\u00b7\u00b7 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . \u00b7\u00b7 12-1 13.0 LAUNCH VEHICLE SUMMARY\u00b7......\u00b7......... 13-1 14.0 ANOMALY SUMMARY \u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . \u00b7\u00b7 14-1 14.2 LUNAR MODULE \u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT \u00b7\u00b7\u00b7 14-36 15.0 CONCLUSIONS\u00b7\u00b7\u00b7\u00b7.\u00b7\u00b7\u00b7.\u00b7\u00b7\uff0e\u00b7.\uff0e\u00b7\u00b7 \u00b7\u00b7 15-1 APPENDIX A - VEHICLE"
  },
  {
    "chunk_id": 15,
    "question": "What is the page number where the seismic detection of third stage lunar impact experiment is discussed?",
    "answer": "11-9",
    "difficulty": "medium",
    "chunk_content": "8.7 TRANSLUNAR FLIGHT ... 8-7 8.8 TRANSEARTH INJECTION 8-11 8.9 TRANSEARTH COAST \u00b7\u00b7\u00b7\u00b7\u00b7 8-11 8.10 ENTRY AND LANDING.\u00b7\u00b7. 8-17 9.0 BIOMEDICAL EVALUATION...... 9-1 9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA . .. 9-1 9.2 INFLIGHT HISTORY \u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7 \u00b7 9-2 9.3 PHYSICAL EXAMINATIONS . .. ? 9-6 10.0 MISSION SUPPORT PERFORMANCE 10-1 10.1 FLIGHT CONTROL \u00b7\u00b7\u00b7\u00b7 10-1 10.2 NETWORK\uff0e\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7 10-2 10.3 RECOVERY OPERATIONS...\u00b7.\u00b7. \u00b7\u00b7\u00b7 10-2 11.0 EXPERIMENTS\u00b7\u00b7\u00b7\u00b7\u00b7\uff0e\u00b7\u00b7\u00b7.\u00b7\u00b7\u00b7\u00b7. \u00b7\u00b7 11-1 11.1 ATMOSPHERIC ELECTRICAL PHENOMENA ....... 11-1 11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES\uff0e\u00b7\uff0e\u00b7\u00b7\uff0e\u00b7\uff0e\u00b7\u00b7\uff0e\u00b7\uff0e\u00b7\u00b7\uff0e\u00b7\u00b7 11-8 11.3 SEISMIC DETECTION OF THIRD STAGE LUNAR IMPACT . 11-9 12.0 ASSESSMENT OF MISSION OBJECTIVES . :. . . . . \u00b7\u00b7 12-1 13.0 LAUNCH VEHICLE SUMMARY\u00b7......\u00b7......... 13-1 14.0 ANOMALY SUMMARY \u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7 14-1 14.1 COMMAND AND SERVICE MODULES . . . . . \u00b7\u00b7 14-1 14.2 LUNAR MODULE \u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7\u00b7 14-24 14.3 GOVERNMENT FURNISHED EQUIPMENT \u00b7\u00b7\u00b7 14-36 15.0 CONCLUSIONS\u00b7\u00b7\u00b7\u00b7.\u00b7\u00b7\u00b7.\u00b7\u00b7\uff0e\u00b7.\uff0e\u00b7\u00b7 \u00b7\u00b7 15-1 APPENDIX A - VEHICLE"
  },
  {
    "chunk_id": 18,
    "question": "What was the time of launch in e.s.t.?",
    "answer": "2:13:00 p.m.",
    "difficulty": "easy",
    "chunk_content": "The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological"
  },
  {
    "chunk_id": 18,
    "question": "Who was substituted as the Command Module Pilot two days before launch?",
    "answer": "John L. Swigert, Jr.",
    "difficulty": "medium",
    "chunk_content": "The space vehicle, with a crew of James A. Lovell, Commander; Fred W. Haise, Jr., Lunar Module Pilot; and John L. Swigert, Jr., Command Module Pilot; was launched from Kennedy Space Center, Florida, at 2:13:00 p.m. e.s.t. (i9:13:00 G.m.t.) April ll, 1970. Two days before launch, the Command Module Pilot, as a member of the Apollo l3 backup crew, was substituted for his prime crew counterpart, who was exposed and found susceptible to rubella (German measles). Prior to launch, a network of meters was installed in the vicinity of the launch site to measure electrical phenomena associated with Saturn V ascent in support of findings from the Apollo l2 lightning investigation; satisfactory data were obtained. During S-II stage boost, an automatic shutdown of the center engine occurred because of a divergent dynamic structural condition associated with that engine. Soon after the spacecraft was ejected, the S-IvB was maneuvered so as to impact on the lunar surface and provide seismological"
  },
  {
    "chunk_id": 20,
    "question": "How long did it take for the pressure in cryogenic oxygen tank 2 to rise at an abnormally high rate?",
    "answer": "Approximately 56 hours",
    "difficulty": "easy",
    "chunk_content": "At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry."
  },
  {
    "chunk_id": 20,
    "question": "What was the reason for the immediate abort of the mission?",
    "answer": "Loss of oxygen and primary power in the service module",
    "difficulty": "medium",
    "chunk_content": "At approximately 56 hours, the pressure in cryogenic oxygen tank 2 began to rise at an abnormally high rate and, within about lo0 seconds, the tank abruptly lost pressure. The pressure in tank l also dropped but at a rate sufficient to maintain fuel cell 2 in operation for approximately 2 more hours. The loss of oxygen and primary power in the service module required an immediate abort of the mission. The crew powered up the lunar module, and the first maneuver following the incident was made with the descent propulsion system to place the spacecraft once again on a free-return trajectory. A second maneuver performed with the descent engine 2 hours after passing pericynthion reduced the transearth transit time and moved the earth landing point from the Indian Ocean to the South Pacific. Two small transearth midcourse corrections were required prior toentry."
  },
  {
    "chunk_id": 22,
    "question": "What section of the report contains a discussion of the abort profile?",
    "answer": "Section 3",
    "difficulty": "easy",
    "chunk_content": "Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7.\n\nA complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified."
  },
  {
    "chunk_id": 22,
    "question": "Where can the current status of all Apollo mission supplements be found?",
    "answer": "Appendix E",
    "difficulty": "easy",
    "chunk_content": "Because an inflight anomaly in the cryogenic oxygen supply required an abort of the mission prior to insertion into lunar orbit, discussions of systems performance only relate to the abort profile and the system configurations required as a result of the emergency. A complete discussion of the anomaly is presented in reference l, and the abort profile is described in section 3. Because of the added criticality of onboard consumables, a discussion of usage profiles in both vehicles is contained in section 7.\n\nA complete analysis of all flight data is not possible within the time allotted for preparation of this report. Therefore, report supplements will be published for certain Apollo l3 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified."
  },
  {
    "chunk_id": 26,
    "question": "What was the location of the resultant landing at earth?",
    "answer": "Indian Ocean",
    "difficulty": "easy",
    "chunk_content": "After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with"
  },
  {
    "chunk_id": 26,
    "question": "How many hours of support did the lunar module systems intend to provide for the crew after the resultant landing?",
    "answer": "90 hours",
    "difficulty": "easy",
    "chunk_content": "After powering up the lunar module, co-aligning the two platforms, and shutting down all command and service module systems following the tank anomaly, a naneuver was immediately performed to return the spacecraft to a free-return profile. The maneuver was performed as the second midcourse correction, using the descent propulsion system in the docked configuration, a mode tested successfully during Apollo 9. The resultant landing at earth would have been at 152 hours in the Indian Ocean, with lunar module systems intended to support the crew for the remaining 90 hours. Because consumables were extremely marginal in this emergency mode and because only minimal recovery support existed at this earth landing location, a transearth injection maneuver using the descent propulsion system was planned for execution 2 hours after passing pericynthion. Between these two maneuvers, an alignment check was made of the lunar module inertial platform to verify the maneuver would be executed with"
  },
  {
    "chunk_id": 28,
    "question": "What was the location of the landing site?",
    "answer": "The South Pacific",
    "difficulty": "easy",
    "chunk_content": "The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control."
  },
  {
    "chunk_id": 28,
    "question": "How many hours before entry was the final midcourse correction performed?",
    "answer": "5",
    "difficulty": "easy",
    "chunk_content": "The transearth injection maneuver was performed on time, and the transearth coast time was shortened such that landing was to occur at about l43 hours in the South Pacific, where primary recovery support was located. Guidance errors during this maneuver necessitated a small midcourse correction at about 105 hours to return the projected entry flight path angle to within specified limits. Following this firing, the spacecraft was maneuvered into a passive thermal control mode, and all lunar module systems were powered down except those absolutely required to support the crew. A final midcourse correction was performed 5 hours before entry to raise the entry flight-path angle slightly, and this maneuver was performed using the lumar module reaction control system under abort guidance control."
  },
  {
    "chunk_id": 30,
    "question": "What was the power source for the three entry batteries?",
    "answer": "Lunar module power",
    "difficulty": "easy",
    "chunk_content": "The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative"
  },
  {
    "chunk_id": 30,
    "question": "Where did the lunar module, including the radioisotope thermoelectric fuel capsule, impact?",
    "answer": "The open sea between Samoa and New Zealand",
    "difficulty": "medium",
    "chunk_content": "The command module was powered up with the three entry batteries, which had been brought up to nearly full charge using lunar module power. The command module platform was aligned to the lunar module platform, and the spacecraft were unaocked 70 minutes before entry. After undocking, the escaping tunnel pressure provided the necessary separation velocity between the two spacecraft. From this point, the mission was completed nominally, as in previous flights, with the spacecraft landing approximately l mile from the target point. The lunar nodule, including the raaioisotope thermoelectric fuel capsule used to power experiment equipment, entered the atnosphere and impacted in the open sea between Samoa and New Zealand at 25.5 degrees south latitude and 176 degrees west longitude, with surveillance aircraft in the area. The three crewmen were onboard the recovery ship, USS Iwo Jima, within 45 minutes of landing, the fastest recovery time for all Apollo manned flights. A narrative"
  },
  {
    "chunk_id": 35,
    "question": "What was the planned orbital altitude of the pericynthion after the first major spacecraft maneuver?",
    "answer": "60 miles",
    "difficulty": "easy",
    "chunk_content": "As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II.\n\nTABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS"
  },
  {
    "chunk_id": 35,
    "question": "What was the achieved pericynthion altitude at translunar injection?",
    "answer": "415.8 miles",
    "difficulty": "medium",
    "chunk_content": "As on Apollo l2, the S-IVB was targeted for a high-pericynthion free-return translunar profile, with the first major spacecraft maneuver intended to lower the pericynthion to the planned orbital altitude of 60 miles. Upon execution of this maneuver, the spacecraft was intentionally placed on a non-free-return trajectory. The achieved pericynthion altitude at translunar injection was 415.8 miles. The accuracy of the translunar injection maneuver was such that the option for the first planned midcourse correction was not exercised. The velocity change required at the second planned midcourse option point, intended as the time for entering the non-free-return profile, was 23.2 ft/sec. The trajectory parameters for the translunar injection and all spacecraft maneuvers are presented in table 4-II.\n\nTABLE 4-I.- DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS"
  },
  {
    "chunk_id": 36,
    "question": "What unit is used to measure altitude above the lunar surface?",
    "answer": "feet or miles",
    "difficulty": "easy",
    "chunk_content": "Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect\uff0cfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial"
  },
  {
    "chunk_id": 36,
    "question": "What is the reference system for the space-fixed velocity vector?",
    "answer": "body-centered, inertial reference coordinate system",
    "difficulty": "medium",
    "chunk_content": "Trafectory Parameters Definition Geodetic latitude Spacecraft position measured north or south from the earth's equator to the local vertical vector, deg Selenographic latitude Spacecraft position measured north or south from the true lunar equatorial plane to the local ver- tical vector, deg Longitude Spacecraft position measured east or west from the body's prime meridian to the local vertical vec- tor,deg Altitude Perpendicular distance from the reference body to the point of orbit intersect\uff0cfeet or miles; alti- tude above the lunar surface is referenced to the altitude of the landing site with respect to mean lunar radius Space-fixed velocity Magmitude of the inertial velocity vector refer- enced to the body-centered, inertial reference coordinate system, ft/sec Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered, local horizontal plane to the inertial velocity vector, deg Space-fixed heading 8ngle Angle of the projection of the inertial"
  },
  {
    "chunk_id": 37,
    "question": "What is the unit of measurement for the maximum altitude above the oblate earth model?",
    "answer": "mile",
    "difficulty": "easy",
    "chunk_content": "velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg"
  },
  {
    "chunk_id": 37,
    "question": "What is the acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane?",
    "answer": "Inclination",
    "difficulty": "medium",
    "chunk_content": "velocity vector onto the local body-centered, horizontal plane, measured positive eastward from north, deg Apogee Maximum altitude above the oblate earth model, mile Perigee Minimum altitude above the oblate earth model, mfle Apocynthi on Maximum altitude above the moon model, referenced to landing Bite altitude, miles Peri cynthi on Minimum altitude above the moon model, referenced to landing site altitude, miles Period Time required for spacecraft to complete 360 de- grees of orbit rotation, min Inclination Acute angle formed at the intersection of the orbit plane and the reference body's equatorial plane, deg Longitude of the ascending node Longitude where the orbit plane crosses the ref- erence body's equatorial plane from below, deg"
  },
  {
    "chunk_id": 39,
    "question": "What was the altitude of the spacecraft above the launch pad at the time of S-IVB second cutoff?",
    "answer": "175.71",
    "difficulty": "easy",
    "chunk_content": "Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N"
  },
  {
    "chunk_id": 39,
    "question": "What was the space-fixed flight-path angle at the time of translunar injection?",
    "answer": "59.318",
    "difficulty": "medium",
    "chunk_content": "Translunar phase Event Reference body Time, hr:min:sec Latitude, aeg Longitude, deg Altitude above launcn : pad, miles Space-fixed velocity, ft/sec Space-fixed fiight-path angle,deg Space-fixed heading angle, deg E of N S-IVB second ignition Earth 2:35:46.4 22.488 142.45E 105.39 25 573.1 .032 65.708 S-IVB second cutoff Earth 2:41:37.2 9.39S 166.45E 175.71 35 562.6 7.182 59.443 Translunar injection Earth 2:41:47.2 8.92S 167.21E 182.45 35 538.4 7.635 59.318 Cormand and service module/S-IVB separation Earth 3:06:38.9 27.03N 129.67W 3 778.54 25 027.8 45.034 72.297 Docking Earth 3:19:08.8 30.21N 118.10W 5 934.90 21 881.4 51.507 79.351 Spacecraft/S-IVB sepa- ration Earth 4:01:00.8 31.95N 105.30W 12 455.83 16 619.0 61.092 91.491 First midcourse correction Ignition Cutofr Second midcourse correction Earth Earth 30:40:49.6 30:40:53.1 22.93N 22.80N 101.85W 101.86w 121 381.93 121 385.43 4 682.5 4 685.6 77.464 77.743 112.843 112.751 Ignition Cutoff Earth Earth 61:29:43.5 61:30:17.7 20.85N 20.74N"
  },
  {
    "chunk_id": 40,
    "question": "What is the time of the Transearth phase?",
    "answer": "79:27 :39.0",
    "difficulty": "easy",
    "chunk_content": "159.70E 159.56E 188 371.38 188 393.19 3 065.8 3 093.2 79.364 79.934 115.464 116.54 Transearth phase Transearth injection Ignition Cutoff Thirdmidcourse correction Moon Moon 79:27 :39.0 79:32:02.8 3.73N 3.62N 65.46E 64.77E 5 465.26 5 658.68 4 547.7 5 020.2 72.645 64.784 -116.308 -117.886 Ignition Earth Earth 105:18:28.0 105:18:42.0 19.63N 19.50N 136.84W 136.90W 152 224.32 152 215.52 4 457.8 4456.6 -79.673 -79.765 114.134 114.242 Fourthmidcourse correction Ignition Cutofr Earth Earth 137:39:51.5 137:40:13.0 11.35N 11.34N 113.39E 113.32E 37 806.58 37 776.05 10 109.1 10 114.6 -72.369 -72.373 116.663 118.660 Service module separation Earth 138:01:48.0 10.88N 108.77E 35 694.93 10405.9 -71.941 118.824 Undocking Earth 141:30:00.2 1.23S 77.55E 11 257.48 1.7 465.9 -60.548 120.621 Entry interface Earth 142:40:45.7 28.23S 173.44E 65.83 36 210.6 -6.269 77.210"
  },
  {
    "chunk_id": 40,
    "question": "What is the latitude of the Earth at the time of the service module separation?",
    "answer": "10.88N",
    "difficulty": "easy",
    "chunk_content": "159.70E 159.56E 188 371.38 188 393.19 3 065.8 3 093.2 79.364 79.934 115.464 116.54 Transearth phase Transearth injection Ignition Cutoff Thirdmidcourse correction Moon Moon 79:27 :39.0 79:32:02.8 3.73N 3.62N 65.46E 64.77E 5 465.26 5 658.68 4 547.7 5 020.2 72.645 64.784 -116.308 -117.886 Ignition Earth Earth 105:18:28.0 105:18:42.0 19.63N 19.50N 136.84W 136.90W 152 224.32 152 215.52 4 457.8 4456.6 -79.673 -79.765 114.134 114.242 Fourthmidcourse correction Ignition Cutofr Earth Earth 137:39:51.5 137:40:13.0 11.35N 11.34N 113.39E 113.32E 37 806.58 37 776.05 10 109.1 10 114.6 -72.369 -72.373 116.663 118.660 Service module separation Earth 138:01:48.0 10.88N 108.77E 35 694.93 10405.9 -71.941 118.824 Undocking Earth 141:30:00.2 1.23S 77.55E 11 257.48 1.7 465.9 -60.548 120.621 Entry interface Earth 142:40:45.7 28.23S 173.44E 65.83 36 210.6 -6.269 77.210"
  },
  {
    "chunk_id": 41,
    "question": "At what time was the S-IVB maneuver to achieve lunar impact initiated?",
    "answer": "6 hours",
    "difficulty": "easy",
    "chunk_content": "The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0."
  },
  {
    "chunk_id": 41,
    "question": "What was the unexplained velocity increase of the S-IVB detected by tracking data at approximately 19 hours 17 minutes?",
    "answer": "5 ft/sec",
    "difficulty": "medium",
    "chunk_content": "The discarded S-IVB stage was targeted for a lunar impact of 3 degrees south latitude and 30 degrees west longitude. The S-IVB maneuver to achieve lunar impact was initiated at 6 hours, with a firing duration of 2l7 seconds using the auxiliary propulsion system. At approximately 19 hours 17 minutes, tracking data indicated the S-IVB had acquired an unexplained velocity increase of about 5 ft/sec along a projected earth radius which altered the projected lunar impact point closer to the target. The stage impacted the lunar surface at 77:56:40 and at a location of 2.4 degrees south latitude and 27.9 degrees west longitude. The targeted impact point was 125 miles from the Apollo l2 seismometer, and the actual point was 74 miles away, well within the desired 189 mile radius. The S-IVB impact results are discussed in section ll.0."
  },
  {
    "chunk_id": 43,
    "question": "What was the time at which the first transearth midcourse correction was performed?",
    "answer": "105:18:28",
    "difficulty": "easy",
    "chunk_content": "The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees .\n\nSpacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12.\n\n(a) Trans lunar"
  },
  {
    "chunk_id": 43,
    "question": "What was the velocity change produced by the transearth midcourse correction?",
    "answer": "7.8 ft/sec",
    "difficulty": "medium",
    "chunk_content": "The first transearth midcourse correction (table 4-III), was performed at 105:18:28 using the descent propulsion system. The firing was conducted at l0 percent throttle and produced a velocity change of about 7.8 ft/sec to successfully raise the entry flight-path angle to minus 6.52 degrees .\n\nSpacecraft navigation for the aborted mission proceeded satisfactorily. Post-pericynthion navigation procedures were designed to support transearth injection, and special data processing procedures were required for dual vehicle tracking prior to entry. Less range data than usual were received from tracking stations during the abort phase because the power amplifier in the spacecraft was turned off for most of the time to conserve electrical power. The small amounts of range data received and the resulting large data arcs, however, were sufficient to maintain navigation accuracies approximately equivalent to those of Apollo 12.\n\n(a) Trans lunar"
  },
  {
    "chunk_id": 44,
    "question": "What was the ignition time of the S-IVB?",
    "answer": "2:35:46.4",
    "difficulty": "easy",
    "chunk_content": "Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57\n\n(b) Transearth"
  },
  {
    "chunk_id": 44,
    "question": "What was the latitude of the pericynthion arrival?",
    "answer": "1.47N",
    "difficulty": "medium",
    "chunk_content": "Maneuver System Ignition time, hr:min:sec Firing time; sec Velocity change. ft/sec Resultant pericynthion conditions Altitude sboveianding site,miles Velocity; ft/sec Latitude, deg Longitude, deg Pericynthion arrival time, hr:min;sec Translunar injection S-IVB 2:35:46.4 350.8 10039 86.8 8184.4 1.47N 178.52E 77:56:22 First midcourse correction Service prcpulsion 30:40:49.6 3.5 23.2 63.2 8277.9 3.34N 178.93E 77:28:39 Second midcourse correction Descentpropulsion 61:29:43.5 34.2 37.8 136. 8053.4 3.02N 179.29W 77:20:57\n\n(b) Transearth"
  },
  {
    "chunk_id": 45,
    "question": "What was the time of the maneuver system ignition?",
    "answer": "79:27:39",
    "difficulty": "easy",
    "chunk_content": "Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46"
  },
  {
    "chunk_id": 45,
    "question": "What was the latitude of the spacecraft at entry arrival time?",
    "answer": "28.22S",
    "difficulty": "medium",
    "chunk_content": "Maneuver System ignitiontime, hr:min:sec Firing time, sec Velocity change, ft/sec Resultant entry interface condition Plight-path angle,deg, Velocity, ft/sec Latitude, deg Longitude, deg Entry arrival time, hr:min:sec Transearth injection Third midcourse correction Descent propulsion 79:27:39 263.6 860.5 No entry (vacuum perigee= 8o.6 miles) Descent prcpulsion 105:18:28 14.0 7.8 -6.24 36 210.6 28.22S 173.49E 142:40:47 Fourth midcourse Lunarmodulereaction correction control 137:39:51.5 21.5 3.0 -6.26 36 210.9 28.23S 173.46E 142:40:46"
  },
  {
    "chunk_id": 46,
    "question": "What was the time of the final midcourse correction maneuver?",
    "answer": "5 hours before entry",
    "difficulty": "easy",
    "chunk_content": "The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel.\n\nThe final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point."
  },
  {
    "chunk_id": 46,
    "question": "Where did the spacecraft land?",
    "answer": "The Pacific Ocean",
    "difficulty": "medium",
    "chunk_content": "The unusual spacecraft configuration required that new procedures for entry be developed and verified. The resulting timeline called for a final midcourse correction 5 hours before entry, separation of the service module 4 hours 39 minutes before entry, and undocking of the lunar module at l hour ll minutes before entry. Service module separation was performed using the lunar module reaction control system. Separation velocity following lunar module umdocking was provided using pressure in the docking tunnel.\n\nThe final midcourse correction maneuver used the lunar module reaction control system. Landing occurred at 142:54:4l in the Pacific Ocean at 2l degrees 38.4 minutes south latitude and 165 degrees 21.7 minutes west longitude, which was about l mile from the target point."
  },
  {
    "chunk_id": 47,
    "question": "What type of propulsion system is discussed in this section?",
    "answer": "Service propulsion",
    "difficulty": "easy",
    "chunk_content": "The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary.\n\n5.1 SITRUCTURAL AND MECHANICAL SYSTEMS"
  },
  {
    "chunk_id": 47,
    "question": "Where are the details of the anomalies discussed?",
    "answer": "In the Anomaly Summary",
    "difficulty": "medium",
    "chunk_content": "The performance of the command and service module systems is discussed in this section. The sequential, pyrotechnic, service propulsion, thermal protection, earth landing, and emergency detection systems and all displays, controls, and crew provisions operated essentially as intended and are not discussed. The pyrotechnic system, which performed all desired functions, did exhibit two minor anomalies, which are discussed only in sections 14.l.6 and 14.l.l0 of the Anomaly Summary, and two discrepancies in the operation of crew equipment were noted, these being discussed in sections 14.3.l and 14.3.2 of the Anomaly Summary. Except for these four cases, all other anomalies are generally mentioned in this Section but are discussed in greater detail in the Anomaly Summary.\n\n5.1 SITRUCTURAL AND MECHANICAL SYSTEMS"
  },
  {
    "chunk_id": 51,
    "question": "How many times were batteries A and B charged during the flight?",
    "answer": "Three",
    "difficulty": "easy",
    "chunk_content": "Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries.\n\n5.2.2 Fuel Cells"
  },
  {
    "chunk_id": 51,
    "question": "What was the source of power for main bus A after the cryogenic oxygen incident?",
    "answer": "Battery A",
    "difficulty": "medium",
    "chunk_content": "Command module battery performance was acceptable throughout the mission. Entry battery C had been isolated throughout the flight, and at 58 hours 40 minutes, batteries A and B were also isolated from the spacecraft buses. Batteries A and B were charged a total of three times each during the flight, including once each using power from the lunar module. Following the cryogenic oxygen incident, battery A was twice placed on main bus A to support spacecraft load requirements. Preentry procedures were conducted with the lunar module supplying power to the command module main bus B through the command and service module/lunar module umbilical and with entry battery C supplying power to main bus A. This configuration was maintained from 6 hours 30 minutes prior to entry until 2 hours 30 minutes prior to entry, at which time the lunar module batteries were disconnected and all electrical power loads were assumed by the command module entry batteries.\n\n5.2.2 Fuel Cells"
  },
  {
    "chunk_id": 52,
    "question": "What was the issue with the fuel cell flow indicators before lift-off?",
    "answer": "Erratic readings",
    "difficulty": "easy",
    "chunk_content": "Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal.\n\nDuring the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l."
  },
  {
    "chunk_id": 52,
    "question": "Which fuel cell continued to operate after the loss of oxygen pressure in tank 2?",
    "answer": "Fuel cell 2",
    "difficulty": "medium",
    "chunk_content": "Prior to lift-off, the crew experienced erratic readings from all three fuel cell flow indicators when cycling the switch, but system operation was normal.\n\nDuring the flight, the three fuel cells operated as expected until the sudden loss of pressure in cryogenic oxygen tank 2, as discussed in Section l4.l.l. Fuel cell 3 condenser exit temperature varied periodically. A behavior present on all previous flights, and characteristic of the system umder certain operating conditions. Soon after the loss of oxygen pressure in tank 2, fuel cells l and 3 lost power and were shut down. Fuel cell 2 sustained the total command and service module load until the depletion of oxygen pressure in tank l."
  },
  {
    "chunk_id": 54,
    "question": "What was the average current of the fuel cells during the mission?",
    "answer": "24 amperes",
    "difficulty": "easy",
    "chunk_content": "During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts.\n\n5.3 CRYOGENIC STORAGE"
  },
  {
    "chunk_id": 54,
    "question": "What was the average bus voltage of the fuel cells during the mission?",
    "answer": "29.4 volts",
    "difficulty": "easy",
    "chunk_content": "During the mission, the fuel cells supplied approximately l20 kW-h of energy at ar average current of approximately 24 amperes per fuel cell and at an average bus voltage of 29.4 volts.\n\n5.3 CRYOGENIC STORAGE"
  },
  {
    "chunk_id": 59,
    "question": "What type of communications were used after translunar injection?",
    "answer": "S-band",
    "difficulty": "easy",
    "chunk_content": "Consumable quantities in the cryogenic storage system are discussed in section 7.l.\n\n5.4 COMMUNICATIONS EQUIPMENT\n\nThe communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good.\n\nPrior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4."
  },
  {
    "chunk_id": 59,
    "question": "How long was the spacecraft powered down?",
    "answer": "approximately 58 hours",
    "difficulty": "medium",
    "chunk_content": "Consumable quantities in the cryogenic storage system are discussed in section 7.l.\n\n5.4 COMMUNICATIONS EQUIPMENT\n\nThe communications system satisfactorily supported the mission. Both S-band and VHF communications were used until translumar injection, after which the VHF was turmed off and the S-band equipment was used until spacecraft power-down at approximately 58 hours. S-band and VHF voice, color television pictures, and real-time and playback telemetry were satisfactory. Uplink and downlink signal strengths corresponded to preflight predictions. Communications system management, including antenna switching, was good.\n\nPrior to the television broadcast at approximately 55 hours , difficulty was experienced with high-gain antenna acquisition for approximately l2 minutes. After a change in spacecraft attitude, satisfactory acquisition was accomplished. Further details concerning this problem are discussed in section 14.l.4."
  },
  {
    "chunk_id": 60,
    "question": "What was the time period during which the communications system was powered up to transmit high-bit-rate telemetry data using the omnidirectional antennas?",
    "answer": "From 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00",
    "difficulty": "medium",
    "chunk_content": "At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data.\n\nFrom 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery.\n\n5.5 INS TRUMENTATION"
  },
  {
    "chunk_id": 60,
    "question": "What was the cause of the damage to the high-gain antenna?",
    "answer": "The loss of a Service module outer panel",
    "difficulty": "easy",
    "chunk_content": "At approximately 56 hours, the high-gain antenna experienced an apparent switch from narrow to wide beamwidth, with a resultant temporary loss of telemetry data. This occurrence coincided with the oxygen tank pressure loss. Post-separation photographs of the service module show damage to the high-gain antenna, which is attributed to the loss of a Service module outer panel. This damage, as discussed in reference l, caused the beam switch and the resultant loss of data.\n\nFrom 101:53:00 to 102:02:00 and from 123:05:00 to 123:12:00, the communications system was powered up to the extent necessary to transmit high-bit-rate telemetry data using the omnidirectional antennas. The S-band system was turned on for verification prior to undocking and performed nominally. The VHF/AM and VHF recovery systems were turned on at parachute deployment and operated nominally throughout recovery.\n\n5.5 INS TRUMENTATION"
  },
  {
    "chunk_id": 61,
    "question": "What was the discrepancy in the suit pressure measurement?",
    "answer": "0.5 psi below cabin pressure",
    "difficulty": "easy",
    "chunk_content": "The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement \u00b7 operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident."
  },
  {
    "chunk_id": 61,
    "question": "Where is the discussion on the potable water quantity measurement anomaly located?",
    "answer": "section 14.l.8",
    "difficulty": "medium",
    "chunk_content": "The instrumentation system performed normally except for the following discrepancies, both of which have occurred on previous flights. The suit pressure measurement indicated 0.5 psi below cabin pressure until the command module was powered down. However, when the command module was powered up at l23 hours, the measurement indicated correct values, as discussed in section l4.i.9. The potable water quantity measurement \u00b7 operated erratically for a brief period early in the mission. This anomaly is described in section 14.l.8. The pressure, temperature, and quantity measurements for oxygen tank 2, along with the fuel cell l nitrogen pressure transducer failure, are discussed in section l4.l.l, since the anomalous performance of these systems is related to the tank incident."
  },
  {
    "chunk_id": 62,
    "question": "What was the reason for the failure of the service propulsion auxiliary propellant gaging system?",
    "answer": "Fuel leakage into the point sensor module within the tank.",
    "difficulty": "easy",
    "chunk_content": "The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected.\n\n5.6 GUIDANCE, NAVIGATION, AND CONTROL"
  },
  {
    "chunk_id": 62,
    "question": "Did the failure of the service propulsion auxiliary propellant gaging system affect the performance of the mission?",
    "answer": "No",
    "difficulty": "easy",
    "chunk_content": "The service propulsion auxiliary propellant gaging system failed prior to launch and a measurement waiver was granted. The failure, which resulted in shorting of the instrumentation power supply, was caused from fuel leakage into the point sensor module within the tank. Similar failures have occurred on previous flights, and since this system is independent of the primary gaging system, which was operating properly, performance of the mission was not affected.\n\n5.6 GUIDANCE, NAVIGATION, AND CONTROL"
  },
  {
    "chunk_id": 65,
    "question": "How long did the initial separation from the S-IvB take?",
    "answer": "4.28 seconds",
    "difficulty": "easy",
    "chunk_content": "All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll."
  },
  {
    "chunk_id": 65,
    "question": "What was the peak rate disturbance in roll during docking?",
    "answer": "0.60 deg/sec",
    "difficulty": "medium",
    "chunk_content": "All attitude control functions were satisfactory. Initial separation from the S-IvB was performed by thrusting for 4.28 seconds to impart a velocity change of 0.86 ft/sec. After a manual pitch maneuver, the command and service modules were docked with the lunar module. Rate disturbances noted at docking were 0.l6 deg/sec peak in pitch and yaw, and 0.60 deg/sec peak in roll."
  },
  {
    "chunk_id": 66,
    "question": "What was the reason for the unsuccessful passive thermal control mode attempt at 7:43:02?",
    "answer": "The roll rate was established using one rather than two roll engines, as required by the checklist.",
    "difficulty": "medium",
    "chunk_content": "The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine"
  },
  {
    "chunk_id": 66,
    "question": "What triggered the unplanned minimum impulse engine firing at 32:21:49?",
    "answer": "The roll manual attitude switch was changed from the rate-command position to the acceleration-command position.",
    "difficulty": "easy",
    "chunk_content": "The passive thermal control modes attempted at 7:43:02 and 32:2l:49 were not successful and had to be reinitiated. The attempt at 7:43:02 resulted in a divergent coning angle because the roll rate was established using one rather than two roll engines, as required by the checklist. In addition, an incorrect roll rate was loaded into the digital autopilot. The attempt at 32:2l:49 resulted in a divergent coning angle because an unplanned minimum impulse engine firing occurred l3 seconds after initiating the roll rate. The engine firing command (two negative roll engines) was generated when the roll manual attitude switch was changed from the rate-command position to the acceleration-command position. The engine"
  },
  {
    "chunk_id": 67,
    "question": "What was the time at which the passive thermal control mode was attempted?",
    "answer": "32:2l:49",
    "difficulty": "easy",
    "chunk_content": "firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally.\n\n\n\nNASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode."
  },
  {
    "chunk_id": 67,
    "question": "What were the adverse effects of the two extraneous firings shown in figure 5.6-1?",
    "answer": "Adverse effects of two extraneous firings",
    "difficulty": "medium",
    "chunk_content": "firing could have been avoided procedurally by disabling all engines before doing any control system switching. The passive thermal control mode attempted at 32:2l:49 is compared with a typical case in figure 5.6-l, which shows the adverse effects of two extraneous firings. All subsequent passive thermal control modes using the command and service module were established normally.\n\n\n\nNASA-5-70-5825 Figure 5.6-l.- Comparison of early translunar maneuver to establish a passive thermal control mode."
  },
  {
    "chunk_id": 68,
    "question": "What caused a computer restart during the oxygen tank incident?",
    "answer": "A voltage transient",
    "difficulty": "easy",
    "chunk_content": "At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly .\n\nThe only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table."
  },
  {
    "chunk_id": 68,
    "question": "How long did it take for the rate and attitude errors to be reduced to a nulled condition?",
    "answer": "75 seconds",
    "difficulty": "easy",
    "chunk_content": "At the time of the oxygen tank incident, three events took place that affected control system performance: the quad C isolation valves closed (as discussed in section l4.l.l), a voltage transient caused a computer restart, and the digital autopilot re-initialized the attitude to which it was referenced. The response of the digital autopilot to these events was as programmed, and rate and attitude errors were reduced to a nulled condition within 75 seconds. Reference l contains a more complete discussion of spacecraft dynamics during and after the oxygen tank anomaly .\n\nThe only translation maneuver performed with the service propulsion system was the first midcourse correction. Spacecraft dynamics during this maneuver were nominal, and sigmificant translation parameters are showm in the following table."
  },
  {
    "chunk_id": 69,
    "question": "What is the time of cutoff for the mission?",
    "answer": "30 : 40 :53.14",
    "difficulty": "easy",
    "chunk_content": "Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1  +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state  Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll\n\nVelocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed."
  },
  {
    "chunk_id": 69,
    "question": "What is the initial pitch of the spacecraft?",
    "answer": "0.95",
    "difficulty": "easy",
    "chunk_content": "Parame ter First midcourse correction Time Ignition, hr:min:sec 30 :40 :49 .65 Cutoff, hr:min:sec 30 : 40 :53.14 3.49 Duration, min:sec Velocity gained, ft/sec* (desirea/actual) X -13.1/-13.2 Y -14.7/-14.5 Z -12.2/-12.3 Velocity residual, ft/sec (spacecraft coordinates)** X +0.1  +0.2 Z +0.3 Entry monitor system +0.7 Engine gimbal. position, deg Initial Pitch 0.95 Yaw -0.19 Maximurn excursion Pitch +0.44 Yaw -0.51 Steady-state  Pitch 1.13 Yaw -0.44 Cutoff Pitch 1.17 M1 -0.44 Maximum rate excursion, deg/sec Pitch +0.08 MB +0.16 Roll -0.08 Maximum attitude error, deg Pitch -0.04 Yaw -0.24 +0.12 Roll\n\nVelocity gained in earth-centered inertial coordinates. **Velocity residuals in spacecraft coordinates after trimning has been completed."
  },
  {
    "chunk_id": 72,
    "question": "What is the time of the first entry in the table?",
    "answer": "00:45",
    "difficulty": "easy",
    "chunk_content": "Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug\uff0c36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263\n\npreferred alignment DRererence metrix (REFSMAT) CCoarse alignment"
  },
  {
    "chunk_id": 72,
    "question": "What is the star used in the coarse alignment at time 40:43?",
    "answer": "36Vega,40Altair",
    "difficulty": "medium",
    "chunk_content": "Time hr:min Optian code Star used Ster angle aifference, deg Gyro torquing angles, deg Gyro drift,mERU Comments X Y Z X Z 00:45 05:26 (a) 26 Spica,33 Antares 0.00 -0.067 -0.000 +0.162 10 : 40 (b) 35 Rasalhague,44 Enif 20 0.01 0.00 +0.175 -0.123 +0.172 -0.012 23:47 (b) Dnoces,27 Alkaid 31 Arcturug\uff0c36 Vega 0.01 -0.283 -0.113 +0.092 28:49 (b) 30 Menkent,32 Alphecca 0.01 -0.084 -0.161 -0.075 +0.403 +1,4 +0.8 +2.1 Check star 36 9:C7 () 23 Denebola,32 Alphecca 0.00 +0.285 +0.011 +0.146 +0.131 +1.1 +1.0 +1.9 Check Btar 35 Check star 31 40 :43 (c) From lunar module primary 0:52 (a) guidance 0.00 36Vega,40Altair ~1.253 +0.385 +3.263\n\npreferred alignment DRererence metrix (REFSMAT) CCoarse alignment"
  },
  {
    "chunk_id": 73,
    "question": "What was the difference in velocity between the S-IvB instrument unit and the command module platform during earth ascent?",
    "answer": "75-ft/sec",
    "difficulty": "easy",
    "chunk_content": "Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift"
  },
  {
    "chunk_id": 73,
    "question": "What was the cause of the discrepancy in the Y-axis error magnitude?",
    "answer": "The magnitude of the null bias drift",
    "difficulty": "medium",
    "chunk_content": "Table 5.6-II summarizes the inertial component preflight histories. Velocity differences between the S-IvB instrument unit and the command module platform during earth ascent indicate a 75-ft/sec difference in the Y-axis. A Y-axis difference is typical of a command module platform gyrocompassing misalignment at lift-off. However, the Y-axis error magnitude is not typical and is the largest observed during ascent to date. The cause of the discrepancy was the magnitude of the null bias drift"
  },
  {
    "chunk_id": 74,
    "question": "What is the X-Scale factor error in ppm?",
    "answer": "2",
    "difficulty": "easy",
    "chunk_content": "Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu\" after ugdut: Accelerometera X-Scule factor error\u3001ppm. 2 -19y 24 7 -199 Bia\uff0ccm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1\uff1f +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis\uff0cmERU/g..\uff0c -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi\u65e5,mERU/g +19.17 7.14 7"
  },
  {
    "chunk_id": 74,
    "question": "What is the acceleration drift in mERU/g for the input axis in the Z-direction?",
    "answer": "-1.0",
    "difficulty": "medium",
    "chunk_content": "Error Sample me an Stendard deviatlon Number of smmple8 Couwtdown valve Flipht load r'lightaverw tloreuflatc lighttvertu\" after ugdut: Accelerometera X-Scule factor error\u3001ppm. 2 -19y 24 7 -199 Bia\uff0ccm/sec -0.18 0.07 1 -0.26 -1 .17 =0.21 =1; , 1f. Y-Scale factor error,ppa. 2 -164 7 -194 -190 Bias,cm/eec -0.20 0.04 7 -0.20 =1.!{ -t). 1? Z-Scale factor error,ppm.: -389 38 7 -419 1 Bias, cm/sec 2 +0.02 0.06 7 +0.07 8_3.0h -i,0: -1).:1.* Cyroacopea X-Rull bias drift,mERU.. Acceleration drift,spin refer- ence axis,mERU/g .. +0.0 -1.:21 1.28 7 +U.5 -.15 Acceleratlon drift,input 0.58 7 -1.0 axis,mERU/g Y-Null biu drirt,mERU. +22.91 -1.34 6.26 7 +s1\uff1f +4. C +1.C -U.04 Acceieration drift,spin refer- 1.88 7 -1.4 ence axis\uff0cmERU/g..\uff0c -0.09 2.05 7 -0.4 +.U Acceleration drirt,input Ax1s,mERU/g +0.11 h.28 7 +l.7 +1.. Z-Null bias drift,mERU. -3.96 1.94 7 -4.0 d_4.9 +1.t9 +v.# Acceleration drift,spin refer-- ence axis,mERU/g.. -5.37 2.56 7 -7.3 -t.0 Acceleration drift,input axi\u65e5,mERU/g +19.17 7.14 7"
  },
  {
    "chunk_id": 77,
    "question": "At what time was power to the guidance and navigation system removed?",
    "answer": "58 hours",
    "difficulty": "easy",
    "chunk_content": "After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\\circ}$ Or $60^{\\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\\mathsf{c m}/\\mathsf{s e c}^{2}$ to the new value of minus $1.66~\\mathsf{c m}/\\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments ."
  },
  {
    "chunk_id": 77,
    "question": "What was the new value of the Z-axis accelerometer bias after the update?",
    "answer": "minus $1.66~\\mathsf{c m}/\\mathsf{s e c}^{2}$",
    "difficulty": "medium",
    "chunk_content": "After the oxygen tank incident, the platform was used as a reference to which the lunar module platform was aligned. All power to the guidance and navigation system, including the inertial measurment unit heaters: was removed at about 58 hours. Heater power was applied about 80 hours later, when the inertial measurement unit was put into standby and the computer turned on. Based upon ground test data and two short periods. of telemetry, the minimum temperature is estimated to have reached $55^{\\circ}$ Or $60^{\\circ}$ F before power-up. The only significant coefficient shift observed after the long cold soak was in the Z-axis accelerometer bias. The shift was compensated for by an update at 14l hours from minus $0.04\\mathsf{c m}/\\mathsf{s e c}^{2}$ to the new value of minus $1.66~\\mathsf{c m}/\\mathsf{s e c}^{2}$ . Although no gyro measurements were obtained just prior to entry, the precision of the landing indicated no large misalignments ."
  },
  {
    "chunk_id": 78,
    "question": "What was the initial accelerometer bias before translunar injection?",
    "answer": "+0.008",
    "difficulty": "easy",
    "chunk_content": "Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.\u00b7 Y Z 96- 116 37 116 Lt- 116 Null bias drift\uff0c mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5\n\nSeveral entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table.\n\n5.7 REACTION CONTROL\n\nTime Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015\n\n5.7.1 Servi ce Module"
  },
  {
    "chunk_id": 78,
    "question": "What was the acceleration drift of the input axis?",
    "answer": "9.0",
    "difficulty": "medium",
    "chunk_content": "Uncompens ated Error term error One-sigma specification Offset velocity, ft/sec X. -0.75 Y Z -0.25 1.19 2 Bias, cm/sec^ X Y Z -0.04 0.2 0.03 0.2 0.099 0.2 Scale factor error, ppm X.\u00b7 Y Z 96- 116 37 116 Lt- 116 Null bias drift\uff0c mERU X. Y Z 2.7 2.0 -0.3 Acceleration drift, input axis mERU/g, 9.0 Acceleration drift, spin reference axis, mERU/g Y. 9.0 5\n\nSeveral entry monitor system bias tests were made during the flight. The associated accelerometer exhibited a stability well within specification limits. Results of each test are given in the following table.\n\n5.7 REACTION CONTROL\n\nTime Time interval . sec Velocity change, ft/sec Accelerometer bias, ft/sec2 Before translunar injection 100 +0.8 +0.008 After translumar injection 100 +1.0 +0.010 10 hours5 minutes 100 +1.8 +0.018 29 hours 40 minutes 100 +1.5 +0.015\n\n5.7.1 Servi ce Module"
  },
  {
    "chunk_id": 79,
    "question": "How much propellant was used for the initial separation from the S-IVB?",
    "answer": "55 pounds",
    "difficulty": "easy",
    "chunk_content": "All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission."
  },
  {
    "chunk_id": 79,
    "question": "What was the difference between predicted and actual propellant usage before the tank anomaly?",
    "answer": "33 pounds",
    "difficulty": "easy",
    "chunk_content": "All service module reaction control parameters were normal from lift-off to the time of the oxygen tank anomaly. A total of 55 pounds of propellant was used for the initial separation from the S-IVB, the turnaround maneuver, docking and ejection. Praor to the tank anomaly, propellant usage was 137 poumds ; 33 pounds less than predicted for that point in the mission."
  },
  {
    "chunk_id": 83,
    "question": "What was the state of the propellant isolation valves during system decontamination at Hawaii?",
    "answer": "Normal (closed)",
    "difficulty": "easy",
    "chunk_content": "System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7.\n\nAll available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation.\n\n5.8 ENVIRONMENIAL CONTROL"
  },
  {
    "chunk_id": 83,
    "question": "What was used to close the system 1 fuel isolation valve?",
    "answer": "Power from ground servicing equipment",
    "difficulty": "easy",
    "chunk_content": "System decontamination at Hawaii was normal, except that the system l fuel isolation valve was foumd to be in the open position. All other propellant isolation valves were in the normal (closed) position. Power from ground servicing equipment was used to close the valve, which operated normally. Postflight investigation of this condition revealed that the electrical lead from the system l fuel-valve closing coil was miswired, making it impossible to apply power to this coil. This anomaly is discussed in section l4.l.7.\n\nAll available flight data and the condition of the system prior to deactivation at Hawaii indicate that the system performed normally from activation through the propellant dump and purge operation.\n\n5.8 ENVIRONMENIAL CONTROL"
  },
  {
    "chunk_id": 85,
    "question": "At what time was the suit compressor turned off?",
    "answer": "56:19:58",
    "difficulty": "easy",
    "chunk_content": "The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power."
  },
  {
    "chunk_id": 85,
    "question": "What was the pressure in the surge tank when it was isolated?",
    "answer": "858 psia",
    "difficulty": "easy",
    "chunk_content": "The oxygen distribution system operated nominally until deactivation following the cryogenic tank incident. The suit compressor was turned off at 56:19:58, and with the repressurization package off line, the surge tank was isolated l7 minutes later at an indicated pressure of 858 psia. The 20-psi system was reactivated briefly four times from the surge tank to pressurize the command module potable water system. Further discussion of oxygen usage is presented in section 7.l. System operation for entry was satisfactory, with the suit compressor limited to a period of operation of only 22 miautes to conserve electri cal power."
  },
  {
    "chunk_id": 92,
    "question": "What was the duration of nominal S-band communications from system actuation?",
    "answer": "Approximately 58 hours",
    "difficulty": "easy",
    "chunk_content": "6.3 COMMUNICATIONS EQUIPMENT\n\nS-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality."
  },
  {
    "chunk_id": 92,
    "question": "What was the primary configuration used for communications?",
    "answer": "Low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband",
    "difficulty": "medium",
    "chunk_content": "6.3 COMMUNICATIONS EQUIPMENT\n\nS-band communications were nominal from system actuation at approximately 58 hours through lunar module undocking. Except for brief periods when high-bit-rate data and high-quality downlink voice were required, low power transmissions, backup voice, and omnidirectional antennas were used to conserve electrical power. The S-band power amplifier was turned off by opening the circuit breaker to provide the higher modulation index for telemetry. The primary cormmunications configuration was low power, low-bit-rate telemetry, omnidirectional antennas, and backup voice on baseband. In this configuration, transmission of high-bit-rate data from the spacecraft was attempted using a 2lo-foot receiving antenna, and except for regular intervals of data dropout because of vehicle attitude charges, these data were of good quality."
  },
  {
    "chunk_id": 93,
    "question": "What was the temperature range of the antenna in the passive thermal control mode?",
    "answer": "plus and minus $25^{\\circ}$ F",
    "difficulty": "easy",
    "chunk_content": "The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66\u00b0 F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\\circ}$ F\u3002\n\n6.4 GUIDANCE, NAVIGATION AND CONTROL\n\nSystem performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit.\n\n6.4.1 Attitude Control"
  },
  {
    "chunk_id": 93,
    "question": "What was the temperature of the antenna when the antenna heaters were turned off?",
    "answer": "minus 66\u00b0 F",
    "difficulty": "easy",
    "chunk_content": "The updata link was used when required and performed nominally. No VHF equipment was exercised, and the S-band steerable antenna was never turned on. The antenna heaters, which normally remain activated, were turned off to conserve power, and the antenna temperature decreased to approximately minus 66\u00b0 F. In the passive thermal control mode, this temperature varied between plus and minus $25^{\\circ}$ F\u3002\n\n6.4 GUIDANCE, NAVIGATION AND CONTROL\n\nSystem performance, with one exception, was nominal during all phases. At completion of the maneuver to the attitude for the last midcourse correction, the attitude error needles were not zeroed because of an out-ofsequence turn-on procedure for the digital autopilot and the inertial measurement urit.\n\n6.4.1 Attitude Control"
  },
  {
    "chunk_id": 100,
    "question": "What was the duration of the first period of throttle increase during the transearth injection maneuver?",
    "answer": "5 seconds",
    "difficulty": "easy",
    "chunk_content": "The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs.\n\n6.4.3 Alignment"
  },
  {
    "chunk_id": 100,
    "question": "What system was used to control yaw during the first transearth midcourse correction?",
    "answer": "abort guidance system attitude-hold mode",
    "difficulty": "medium",
    "chunk_content": "The transearth injection maneuver was performed with the primary guidance system controlling the descent propulsion system. The throttle profile was 5 seconds at 12.6 percent, 2l seconds at 40 percent, and the remainder at full throttle. During both periods of throttle increase, the roll-gimbal drive actuator traveled approximately l.35 degrees negatively from its value at ignition. These excursion were somewhat larger than expected, but simulations have since shown them to be normal and result from engine compliance and mistrim. Spacecraft dynamics were nominal throughout the firing. The first transearth midcourse correction was the last maneuver to use the descent propulsion system. The maneuver was performed by manually controlling pitch and roll using the hand controllers and by automatically controlling yaw with the abort guidance system attitude-hold mode. The l4-second firing was accomplished at 10-percent throttle with no adverse dynami cs.\n\n6.4.3 Alignment"
  },
  {
    "chunk_id": 102,
    "question": "What was the duration of the third midcourse correction?",
    "answer": "34.23",
    "difficulty": "easy",
    "chunk_content": "Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# \u4eba +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44         Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 \u00b10.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg  Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60"
  },
  {
    "chunk_id": 102,
    "question": "What was the maximum rate excursion of the pitch gimbal drive actuator in degrees per second?",
    "answer": "-0.6",
    "difficulty": "medium",
    "chunk_content": "Condition Maneuver Second midcourse correction Transearth injection Third midcourse correction Fourth midcourBe correction .PGNCS/DPS PGNCS/DPS AGS/DPS AGS/DPS Time Ignition,hr:min:sec Cutoff,hr:min:sec Duration,sec 61:29:43.49 61:30:17.72 34.23 79:27:38.95 79 :32:02.77 105:18:28 105:18:42 137:39:51.5 137:40:13 Velocity change before trin (actual/desired) X# \u4eba +3.0/+2.9 -34.2/-34.3 -15.9/-16.2 -425.9/-426.4 +644.6/+645.6 +378.8/+379.0 7.6/7.8 -1.2/-1.5 -1.9/-2.2 Velocity residual after trim, ft/sec X Y Z +0.2 0.0 +0.3 +1.0 +0.3 0.0 ## 0 0.1 Gimbal drive actuator, in, Initial Pitch Roll Maximum excursion Pitch Roll -0.02 -0.34 +0.31 +0.13 -0.28 Not applicable Not applicable Steady-state -0.27 +0.16 -0.44         Pitch +0.04 +0.21 Roll -0.51 -0.55 Cutoff Pitch +0.10 +0.23 Rol1 -0.31 -0.55 Maximum rate excursion,deg/sec Pitch -0.6 +0.2 Roll -0.8 +0.2 +0.2 +0.8 -0.6 +0.2 \u00b10.2 +0.4 +0.2 +0.2 Maximum attitude excursion, deg  Pitch -3.62 -1.6 -0.6 -0.4 Rol1 +1.69 +6.7 +0.9 -0.6 Yaw -1.60"
  },
  {
    "chunk_id": 105,
    "question": "What was the standard deviation of the scale factor error for the X-axis accelerometer?",
    "answer": "-681",
    "difficulty": "easy",
    "chunk_content": "The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table.\n\nSample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis\uff0cmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.\u00b7 -5.38 2.37 4 -5.5 -4.0\n\n6.4.5 Abort Guidance System Performance"
  },
  {
    "chunk_id": 105,
    "question": "How many samples were used to calculate the average bias for the Y-axis accelerometer?",
    "answer": "5",
    "difficulty": "easy",
    "chunk_content": "The inertial measurement unit performed properly throughout the mission. A preflight history of the inertial components and the inflight accelerometer bias measurements are given in the following table.\n\nSample me an St andard deviation Number of samples Countdown value Flight load Flight average Accelerometers X - Scale factor error,ppm -681 Bias,cm/sec 2 +1.47 Y - Scale factor error, ppm -1165 5 0.06 18 0.065 4 4 4 ~689 +1.4 ~1173 -1.42 -700 +1.49 -1190 -1.42 -310 +1.50 -1.35 Z - Scale factor error, ppm -244 61 4 Bias, cm/sec 2 +1.56 0.017 4 +1.57 +1.56 +1.52 X- Null bias drift,mERU. +1.18 1.33 4 +0.2 +0.4 Acceleration drift, spin refer- ence axis\uff0cmERU/g. -0.93 1.19 4 -2.6 -1.0 Acceleration drift, input axis, mERU/g.\u00b7 -5.38 2.37 4 -5.5 -4.0\n\n6.4.5 Abort Guidance System Performance"
  },
  {
    "chunk_id": 107,
    "question": "What is the standard deviation of the accelerometer bias in the X direction?",
    "answer": "16.3",
    "difficulty": "easy",
    "chunk_content": "Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 \u4eba -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee\uff0c deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 \u4eba -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr\n\n6.5 REACTION CONTROL"
  },
  {
    "chunk_id": 107,
    "question": "What is the mean gyro spin axis mass in the X direction?",
    "answer": "0.86",
    "difficulty": "medium",
    "chunk_content": "Accelerometer bias Sample me an, Standard deviation, Number of Bample8 Final cali- bration value, Fiight load, X 36.9 16.3 18 57.0 60.0 \u4eba -32.6 10.0 18 -32.0 -31.0 2 -1.6 32.3 18 16.0 47.0 Accelerometer scale factor Standard deviation, udd Number of 8amples Final cali- braticn value, dd Flight lcad, udd X 15.0 18 266 266 Y 16.0 18 -1222 -1249 Z 14.0 18 -&05 -822 Gyroscalefactor Sample meen, ppm St andard deviation, udd Number of samples Final cali- bration value, udd Flight lo8d, PPm X 895 8.7 18 899 898 Y 863 12.9 18 870 870 2 1495 9.5 18 1501 1502 Gyro fixed drift Sample mee\uff0c deg/hr Standard deviation, Number or Final cali- bration value, Flight load, X 0.02 deg/hr 0.08 Samples 18 deg/hr 0.11 deg/hr 0.06 \u4eba -0.30 0.06 18 -0.29 Z -0.58 0.06 18 -0.45 -0.30 Gyro spin axis mass Sample Standard Number Final cali- -0.47 X mean deg/hr 0.86 deviatlon, deg/hr 0.10 or gamples 18 bration value, deg/hr 0.90 Might load, deg/nr\n\n6.5 REACTION CONTROL"
  },
  {
    "chunk_id": 108,
    "question": "What was the initial propellant consumption of the reaction control system?",
    "answer": "467 pounds",
    "difficulty": "easy",
    "chunk_content": "The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds.\n\nAbout 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open.\n\nDuring passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\\circ}$ to97\u00b0F\u3002\n\n6.6 DESCENT PROPULSION"
  },
  {
    "chunk_id": 108,
    "question": "What was the duration of the decrease in system-A propellant manifold pressures?",
    "answer": "4 or 5 seconds",
    "difficulty": "medium",
    "chunk_content": "The reaction control system was activated at about 58 hours. Total propellant consumption was 467 pounds.\n\nAbout 6 minutes after activation, flight data showed a sizeable decrease (approximately 22 psi) in the system-A propellant manifold pressures. This decrease continued for about 4 or 5 seconds and was accompanied by an increase of 7 and 8 psi in the ascent propulsion system fuel and oxidizer manifold pressures, respectively. These manifold pressure changes indicate a high flow rate from the reaction control system. This was verified by a decrease in the indicated quantity by about 15 pounds At this same time, the indicated position for the system-A ascent-feed interconnect valves was open.\n\nDuring passive thermal control modes, the cluster heaters were not used and cluster temperatures ranged from $55^{\\circ}$ to97\u00b0F\u3002\n\n6.6 DESCENT PROPULSION"
  },
  {
    "chunk_id": 110,
    "question": "What was the duration of the transearth injection maneuver?",
    "answer": "264 seconds",
    "difficulty": "easy",
    "chunk_content": "The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained."
  },
  {
    "chunk_id": 110,
    "question": "At what throttle position was the pressurization isolation Solenoid closed?",
    "answer": "l2 percent of full thrust",
    "difficulty": "medium",
    "chunk_content": "The descent propulsion system performed normally during the 34.3- second midcourse correction to enter a free-return trajectory. This maneuver was begun at the minimum throttle position (l2 percent of full thrust), and after 5 seconds, the throttle position was manually increased to approximately 37 percent, which was maintained for the remainder of the firing. The transearth injection maneuver lasted 264 seconds. Approximately 15 seconds prior to engine shutdowm, the pressurization isolation Solenoid was closed to avoid a possible problem with propellant-tank fracture mechanics, and the maneuver was completed in the blowdown mode in which residual helium is the sole pressure source. The third system firing, a midcourse correction maneuver, was l3.7 seconds in duration and was performed in the blowdown mode at the minimum throttle position. Upon completion of this third and final descent propulsion operation, more than half the initial propellant load remained."
  },
  {
    "chunk_id": 111,
    "question": "What was the nominal ground pressure rise rate during the countdown demonstration test?",
    "answer": "7.8 psi/hr",
    "difficulty": "easy",
    "chunk_content": "The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates.\n\nThe average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l."
  },
  {
    "chunk_id": 111,
    "question": "What was the average rise rate from lift-off to the first descent propulsion maneuver?",
    "answer": "7.0 psi/hr",
    "difficulty": "easy",
    "chunk_content": "The supercritical helium pressurization system displayed abnormal performance, beginning with preflight operations. Prelaunch measurements taken during the countdown demonstration test indicated a nominal ground pressure rise rate of 7.8 psi/hr. However, other special tests were performed at various conditions which gave significantly higher rise rates.\n\nThe average rise rate from lift-off to the first descent propulsion maneuver was 7.0 psi/hr. Between the first and second firings, the rise rate increased to l0.5 psi/hr, and after the second firing, the rise rate was 33.5 psi/hr. This anomaly is further discussed in section 14.2.l."
  },
  {
    "chunk_id": 112,
    "question": "What was the predicted rupture range for the vehicle's helium tank?",
    "answer": "190o \u00b1 20 psia",
    "difficulty": "easy",
    "chunk_content": "At about l09 hours when the helium bottle pressure had reached approximately l937 psi, the burst diaphragm ruptured and relieved the supercritical system through a special non-propulsive vent. The predicted rupture range for this vehicle was 190o \u00b1 20 psia. During venting, unexpected motion was inparted to the spacecraft which disrupted the motion established for the passive thermal control mode. The vent tube for the supercritical helium tank is ported on two sides by diametrically opposed Oval-shaped holes. It was originally believed that the escaping gas would exit these holes at 9o degrees to the tube axis such that no net thrust is produced. However, the pressure distribution in the tube is such that the two gas plumes have an included angle less than 180 degrees and probably closer to 9o degrees. Therefore, the component of the gas flow along the axis of the vent tube produces a net thrust in the opposite direction which tends to induce a slight roll rate to the vehicle. Since"
  },
  {
    "chunk_id": 112,
    "question": "What shape are the holes ported on two sides of the vent tube for the supercritical helium tank?",
    "answer": "Oval-shaped",
    "difficulty": "medium",
    "chunk_content": "At about l09 hours when the helium bottle pressure had reached approximately l937 psi, the burst diaphragm ruptured and relieved the supercritical system through a special non-propulsive vent. The predicted rupture range for this vehicle was 190o \u00b1 20 psia. During venting, unexpected motion was inparted to the spacecraft which disrupted the motion established for the passive thermal control mode. The vent tube for the supercritical helium tank is ported on two sides by diametrically opposed Oval-shaped holes. It was originally believed that the escaping gas would exit these holes at 9o degrees to the tube axis such that no net thrust is produced. However, the pressure distribution in the tube is such that the two gas plumes have an included angle less than 180 degrees and probably closer to 9o degrees. Therefore, the component of the gas flow along the axis of the vent tube produces a net thrust in the opposite direction which tends to induce a slight roll rate to the vehicle. Since"
  },
  {
    "chunk_id": 114,
    "question": "How many hours did the environmental control system provide a habitable environment for the crew?",
    "answer": "83",
    "difficulty": "easy",
    "chunk_content": "6.7 ENVIRONMENTAL CONTROL\n\nEnvironmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5.\n\nAn indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr."
  },
  {
    "chunk_id": 114,
    "question": "What was the average water usage rate at the higher end of the range?",
    "answer": "6.3 lb/hr",
    "difficulty": "easy",
    "chunk_content": "6.7 ENVIRONMENTAL CONTROL\n\nEnvironmental. control system performance was satisfactory during the abort phase of the mission and provided a habitable environment for the crew for approximately 83 hours, nearly twice the time of a nominal flight. Only one anomaiy, reverse oxygen leakage through one of the as cent stage shutoff valves, occurred but did not compromise system performance. All crew provisions performed as intended except for cracking of a window shade , discussed in 14 .2.5.\n\nAn indicated total of approximately 290 pounds of water was used from the lumar module tanks between activation of the sublimator and undocking, and an indicated total of about 50 pounds of water remained. Most of the water used for drinking and food preparation was obtained from the command rnodule potable water tank before $\\mathtt{124}$ hours , and drinking water was subsequently used from the lunar module tanks. Average water usage rates varied between 2.6 and 6.3 lb/hr."
  },
  {
    "chunk_id": 117,
    "question": "What was partially restricted with tape in the suit-circuit arrangement?",
    "answer": "The mass flow",
    "difficulty": "easy",
    "chunk_content": "in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft."
  },
  {
    "chunk_id": 117,
    "question": "How long was the lunar module suit circuit operating before an additional unit was stacked on each original cartridge?",
    "answer": "Approximately 20 hours",
    "difficulty": "medium",
    "chunk_content": "in figure 6.7-l, one corner of the bag was Sealed to the inlet of the suit circuit hose. The cabin atmosphere then returned to the lunar module suit circuit through these supplemental cartridges by way of the two outlet hoses. The mass flow through this arrangement was partially restricted with tape to properly load the suit-circuit compressors. After approximately 20 hours of operation with two command module cartridges, an additional unit was stacked on each original cartridge to improve the carbon dioxide removal capability. With this supplemental configuration, when only command module cartridges were being used, the indicated carbon dioxide level was maintained between 0.l arid i.8 mm Hg. The supplemental removal configuration using the command module lithium hydroxide cartridge was assembled and tested on the ground during the flight prior to its actual use in the spacecraft."
  },
  {
    "chunk_id": 121,
    "question": "What was the total amount of propellant loaded into the tanks for the service propulsion system?",
    "answer": "25084",
    "difficulty": "easy",
    "chunk_content": "Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs.\n\n7.l.l Service Propulsion Propellants\n\nThe service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off.\n\nFuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines \"79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7\n\n7.1.2 Reaction Control Propellants"
  },
  {
    "chunk_id": 121,
    "question": "What was the amount of fuel consumed by the service propulsion system?",
    "answer": "92.3",
    "difficulty": "easy",
    "chunk_content": "Consumable usages for the command and service modules prior to the incident were nominal. Following the incident and the attendant shutdowm of command module power, the only consumables used prior to entry were drinking water and surge-tank oxygen, required to pressurize the potable water tank. Specific consumable usages for appropriate systems are presented in the following paragraphs.\n\n7.l.l Service Propulsion Propellants\n\nThe service propulsion system was used only for the first midcourse correction. The propellant loadings listed in the following table were calculated from gaging system readings and measured densities prior to lift-off.\n\nFuel, 1b Oxi di zer, lb Total Loaded In tanks 15 606 24960 In lines \"79 124 Total 15 685 25084 40769 Consume d 92.3 147 239.3 Remaining at time of incident 15 592.7 24937 40 529.7\n\n7.1.2 Reaction Control Propellants"
  },
  {
    "chunk_id": 124,
    "question": "What was the total oxygen supply in the surge tank after the incident?",
    "answer": "3.77 pounds",
    "difficulty": "easy",
    "chunk_content": "Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0\n\n7.1.4 Oxy ge n\n\nFollowing the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry.\n\n7.1.5 Water\n\nAt the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation.\n\n7.1.6 Batteries"
  },
  {
    "chunk_id": 124,
    "question": "How much water was transferred from the command module to the lunar module during the abort phase?",
    "answer": "14 pounds",
    "difficulty": "easy",
    "chunk_content": "Hy drogen, l.b Oxygen,1b Available at lift-off Tank 1 29.0 326.8 Tank 2 29.2 327.2 Totals 58.2 654.0 Con sume d Tank 1 7.1 71.8 Tank 2 6.9 85.2 Totals 14.0 157.0 Remaining at the time of the ncident Tank 1. 21.9 255.0 Tank 2 22.3 242.0 Tot als 44.2 497.0\n\n7.1.4 Oxy ge n\n\nFollowing the incident and loss of pressure in tank l, the total oxygen supply consisted of 3.77 pounds in the surge tank and l pound in each of the three repressurization bottles. About 0.6 pound of the oxygen from the surge tank was used during potable water tank pressurizations and to activate the oxygen system prior to entry. An additional 0.3 pound was used for breathing during entry.\n\n7.1.5 Water\n\nAt the time of the incident, about 38 pounds of water was available in the potable water tank. During the abort phase, the crew used juice bags to transfer approximately 14 pounds of water from the command module to the lunar module for drinking and food preparation.\n\n7.1.6 Batteries"
  },
  {
    "chunk_id": 125,
    "question": "What was the amount of energy remaining in the batteries at landing?",
    "answer": "29 ampere-hours",
    "difficulty": "easy",
    "chunk_content": "The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained.\n\n\n\nFollowing lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident.\n\n7.2.l Des cent Propulsion Propellants\n\nThe loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off."
  },
  {
    "chunk_id": 125,
    "question": "What was the original amount of energy remaining in the three entry batteries?",
    "answer": "99 ampere-hours",
    "difficulty": "easy",
    "chunk_content": "The command module was completely powered down at 58 hours 40 minutes, at which time 99 ampere-hours remained in the three entry batteries. By charging the batteries with lunar module power, available battery capacity was increased to ll8 ampere-hours. Figure 7.l-l depicts the battery energy available and used during entry. At landing, 29 ampere-hours of energy remained.\n\n\n\nFollowing lunar module power-up, oxygen, water, and battery power were consumed at the lowest practical rate to increase the duration of spacecraft support from a nominal 44 hours to a required 83 hours plus margins. In addition, the descent propulsion and reaction control systems were used to effect all required translation and attitude maneuvers following the incident.\n\n7.2.l Des cent Propulsion Propellants\n\nThe loaded quantities of descent propulsion system propellants shown in the following table were calculated from quantity readings in the spacecraft and measured densities prior to lift-off."
  },
  {
    "chunk_id": 126,
    "question": "What was the total amount of fuel loaded in System A?",
    "answer": "316.5",
    "difficulty": "easy",
    "chunk_content": "Fuel, 1b Oxi di zer\uff0c lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6\n\n7.2.2 Reaction Control Propellants\n\nThe reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature.\n\nFuel, lb Oxidi zer\uff0c lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5  Total 166\n\n7.2.3 0xygen\n\nActual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data.\n\nLoaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4"
  },
  {
    "chunk_id": 126,
    "question": "How much oxygen was consumed by the Ascent stage Tank 2?",
    "answer": "2.4",
    "difficulty": "medium",
    "chunk_content": "Fuel, 1b Oxi di zer\uff0c lb Total Loaded 7083.6 11 350.9 18 434.5 Consumed 3225.5 5117.4 8 342.9 Remaining at undocking 3858.1 6 233.5 10 091.6\n\n7.2.2 Reaction Control Propellants\n\nThe reaction control system propellant consumption, shown in the following table, was calculated from telemetered helium tank data using the relationship between pressure , volume, and temperature.\n\nFuel, lb Oxidi zer\uff0c lb Total Loaded System A 107.7 107.7 208.8 208.8 316.5 316.5 ' System B 633.0 Total Consumed System A System B 220 247 Total 467 Remaining at undocking System A 96.5 System B 69.5  Total 166\n\n7.2.3 0xygen\n\nActual oxygen usage closely followed predicted rates from the time of lunar module power-up until undocking, at which time approximately 32 pounds of oxygen remained. The values in the following table are based on telemetered data.\n\nLoaded; 1b Consumed, 1b Remaining after undocking, lb Descent stage 49.3 21.9 27.4 Ascent stage Tank 1 2.3 2.3 Tank 2 2.4 82.7 Total 54.0 21.9 32.4"
  },
  {
    "chunk_id": 127,
    "question": "What was the total loaded water available for inflight use at launch?",
    "answer": "338 pounds",
    "difficulty": "easy",
    "chunk_content": "arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4).\n\n7.2.4 Water\n\nDuring the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking\u00b7is shown in figure 7.2-l.\n\n7.2.5 Batteries"
  },
  {
    "chunk_id": 127,
    "question": "How many hours of cooling would 50 pounds of water provide at the reduced power condition?",
    "answer": "18 hours",
    "difficulty": "medium",
    "chunk_content": "arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4).\n\n7.2.4 Water\n\nDuring the abort phase, lunar module water, which is used primarily to cool the cabin and onboard equipment, was the most restrictive consumable. As a result, extreme measures were taken to shut down all nonessential equipment in order to provide the maximum margin possible. At launch, the total loaded water available for inflight use was 338 pounds. At the time of undocking, approximately 50 pounds of water remained and, at the reduced power condition, would have provided an additional 18 hours of cooling. The actual water usage from the time of initial power-up to undocking\u00b7is shown in figure 7.2-l.\n\n7.2.5 Batteries"
  },
  {
    "chunk_id": 128,
    "question": "What was the initial current consumption of the vehicle before the second descent propulsion system firing?",
    "answer": "30 amperes",
    "difficulty": "easy",
    "chunk_content": "At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to\n\noperate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking.\n\n\n\n\n\nFigure 7.2-2.- Lunar module total battery capacity during flight.\n\n\n\nApollo 13 flight crew"
  },
  {
    "chunk_id": 128,
    "question": "How many ampere hours remained in the lunar module batteries at the time of undocking?",
    "answer": "410",
    "difficulty": "easy",
    "chunk_content": "At the time of power up, 2l79 ampere-hours of electrical energy was available from the four descent- and two ascent-stage batteries. As indicated in figure 7.2-2, initial consumption was at a current of 30 amperes until the second descent propulsion system firing, after which the vehicle was powered down to a l2-ampere load. At approximately ll2 hours, power . was provided to charge the command module entry batteries at a rate of about 7 amperes for approximately l5 hours. The command module was also powered from the lunar module at an ll-ampere rate for a brief period to\n\noperate the reaction control heaters and telemetry equipment. The estimated total energy transferred to the command module was approximately 129 ampere hours. A total of 410 ampere hours remained in the lunar module batteries at the time of undocking.\n\n\n\n\n\nFigure 7.2-2.- Lunar module total battery capacity during flight.\n\n\n\nApollo 13 flight crew"
  },
  {
    "chunk_id": 129,
    "question": "What was the name of the Lunar Module Pilot mentioned in the text?",
    "answer": "Fred W. Haise, Jr.",
    "difficulty": "easy",
    "chunk_content": "Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr.\n\n8.0 PIIOTS' REPORT\n\n8.1 TRAINING"
  },
  {
    "chunk_id": 129,
    "question": "Who was the Command Module Pilot mentioned in the text?",
    "answer": "John L. Swigert, Jr.",
    "difficulty": "easy",
    "chunk_content": "Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr.\n\n8.0 PIIOTS' REPORT\n\n8.1 TRAINING"
  },
  {
    "chunk_id": 130,
    "question": "What was the location of the crew's training base until December 1969?",
    "answer": "Houston",
    "difficulty": "easy",
    "chunk_content": "Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as \"observers\" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive"
  },
  {
    "chunk_id": 130,
    "question": "What was a primary objective of the training for the Apollo 13 mission?",
    "answer": "A field geology experiment as part of the second extravehicular excursion",
    "difficulty": "medium",
    "chunk_content": "Crew training for Apollo 13 commenced on August l, 1969. The crew was based in Houston until December l, l969, when operations were transferred to the launch site.for final training. The training time was adequate to meet the planned launch date of April ll, 1970, and all training objectives were met. The only difficulty in coordinating the training activities was the scheduling of the lunar landing training vehicle for the Commander. The late availability of this vehicle, the large amount of time required for this type of training, and the need to travel between Houston and Cape Kennedy complicated the training Schedule significantly. Because a primary objective was a field geology experiment as part of the second extravehicular excursion, considerable emphasis was placed on geology training. A week-long geology field trip to train the crew as \"observers\" was completed early in the training cycle. Later field trips emphasized practical geological procedures and timelines. Extensive"
  },
  {
    "chunk_id": 134,
    "question": "What was the time at which S-II staging and S-IVB ignition occurred?",
    "answer": "9 minutes 57 seconds",
    "difficulty": "easy",
    "chunk_content": "Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three"
  },
  {
    "chunk_id": 134,
    "question": "How long after S-II staging did engine 5 shut down?",
    "answer": "2 minutes",
    "difficulty": "medium",
    "chunk_content": "Ignition and lift-off occurred on schedule. A listing_of major flight plan events as they occurred is contained in figure 8-l. Firststage performance was nominal and coincided very closely with simulations. Communications during the high noise level phase of flight were excellent. Staging of the S-IC occurred nearly on time and was accompanied by three distinct longitudinal oscillations. S-Il ignition and thrusting was smooth until about 00:05:32, when a sudden buildup in vibration was felt, followed by illumination of the number 5 engine out light. The Mission Control Center confirmed that engine 5 had shut down approximately 2 minutes early. S-II performance after that time was smooth with no noticeable abnormalities. S-II staging and S-IVB ignition occurred late, at 9 minutes 57 seconds. S-IVB performance was nominal but seemed to be accompanied by more vibration than was noted during Apollo 8. [The Apollo l3 Commander had been the Command Module Pilot for Apollo 8]. All three"
  },
  {
    "chunk_id": 135,
    "question": "What was the time of S-IVB engine cutoff?",
    "answer": "00:12:30",
    "difficulty": "easy",
    "chunk_content": "crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles."
  },
  {
    "chunk_id": 135,
    "question": "What was the velocity of the spacecraft at insertion?",
    "answer": "25 565 ft/sec",
    "difficulty": "medium",
    "chunk_content": "crewmen noted the small change in acceleration caused by the mixture ratio shifts during S-II and S-IVB flight. S-IVB engine cutoff occurred at 00:12:30, with the spacecraft guidance system registering the following insertion parameters: velocity 25 565 ft/sec, apogee 102.6 miles, and perigee l00.l miles."
  },
  {
    "chunk_id": 137,
    "question": "What was the range of hours during which the S-IVB vibration was experienced?",
    "answer": "69 to 122 hours",
    "difficulty": "easy",
    "chunk_content": "Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency ,\n\n\n\nFigure 8-l.- Flight plan activities.\n\n\n\n\n\nFigure 8-l.- Continued\n\n\n\n(c) 69 to 122 hours. Figure 8-l.- Continued.\n\n\n\n(a) 122 to 143 hours. Figure 8-l.- Concluded.\n\nlow-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value.\n\n8.6 TRANSPOSITION AND DOCKING"
  },
  {
    "chunk_id": 137,
    "question": "What was the computer-displayed inertial velocity at cutoff?",
    "answer": "35 560 ft/sec",
    "difficulty": "medium",
    "chunk_content": "Nominal first-opportunity translunar injection procedures were used and are satisfactory. Based on S-IVB orbit attitude hold, the ground controllers updated the spacecraft attitude indicators from 18 to 20 degrees. This update was satisfactory and resulted in an essentially zero theta angle in the orbital rate display during the S-IVB translunar injection. S-IVB vibration was greater during translunar injection than that experienced during Apollo 8. These vibrations had high-frequency ,\n\n\n\nFigure 8-l.- Flight plan activities.\n\n\n\n\n\nFigure 8-l.- Continued\n\n\n\n(c) 69 to 122 hours. Figure 8-l.- Continued.\n\n\n\n(a) 122 to 143 hours. Figure 8-l.- Concluded.\n\nlow-magnitude characteristics but presented no problems for monitoring of the injection maneuver. At cutoff, the computer-displayed inertial velocity was 35 560 ft/sec, and the entry monitor system accelerometer confirmed the maneuver to be within 3 ft/sec of the desired value.\n\n8.6 TRANSPOSITION AND DOCKING"
  },
  {
    "chunk_id": 138,
    "question": "What was the rate of the manual pitch maneuver executed after separation and translation?",
    "answer": "1.5 deg/sec",
    "difficulty": "easy",
    "chunk_content": "Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds\n\n8.7 TRANSLUNAR FLIGHT\n\n8.7.1 Coast Phase Activities"
  },
  {
    "chunk_id": 138,
    "question": "How much reaction control fuel was used for transposition, docking, and extraction?",
    "answer": "55 pounds",
    "difficulty": "easy",
    "chunk_content": "Following separation and translation, a manual pitch maneuver of 1.5 deg/sec was executed. Computer control was reselected, and a translation was initiated to give a small closing velocity. A digital autopilot maneuver was executed to aligm the respective roll attitudes. Maximum spacecraft separation was approximately 80 feet. At the final attitude, the image in the crewman optical alignment sight was almost completely washed out by the sun reflection from the lunar module until the vehicles were separated by 6 feet or less. Contact was made at approximately 0.2 ft/sec with a slight roll misalignment. Subsequent tunnel inspection revealed a roll index angle of minus 2.0 degrees. The handles on latches l and 4 were not locked and were recocked and released manually. Spacecraft ejection was normal. Total reaction control fuel used for transposition, docking, and extraction was reported as 55 pounds\n\n8.7 TRANSLUNAR FLIGHT\n\n8.7.1 Coast Phase Activities"
  },
  {
    "chunk_id": 139,
    "question": "How long did earth weather photography last?",
    "answer": "approximately 6 hours",
    "difficulty": "easy",
    "chunk_content": "Following translunar injection, earth weather photography was conducted for approximately 6 hours.\n\nThe first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget."
  },
  {
    "chunk_id": 139,
    "question": "What was the goal of the first period of translunar navigation?",
    "answer": "to establish the apparent horizon attitude for optical marks in the computer",
    "difficulty": "medium",
    "chunk_content": "Following translunar injection, earth weather photography was conducted for approximately 6 hours.\n\nThe first period of translunar navigation (Program 23) at 6 hours was done to establish the apparent horizon attitude for optical marks in the computer. Some manual maneuvering was required to achieve a parallel reticle pattern at the point of horizon-star superposition. The second period of navigation measurements was less difficult, and both periods were accomplished within the timeline and reaction control fuel budget."
  },
  {
    "chunk_id": 140,
    "question": "What roll rate was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole?",
    "answer": "0.3 deg/sec",
    "difficulty": "easy",
    "chunk_content": "The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform\n\nalignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed.\n\nAt about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground.\n\n8.7.2 First Midcourse Correction"
  },
  {
    "chunk_id": 140,
    "question": "What condition was confirmed by the ground at about 47 hours?",
    "answer": "Oxygen tank 2 quantity sensor failed full scale high",
    "difficulty": "medium",
    "chunk_content": "The passive thermal control mode was initiated with the digital autopilot. A roll rate of 0.3 deg/sec was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole. An incorrect entry procedure was used on one attempt and reinitialization of passive thermal control was required. After proper initialization, all thrusters were disabled and the spacecraft maintained an attitude for thermal protection for long periods without approaching gimbal lock. Platform\n\nalignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed.\n\nAt about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground.\n\n8.7.2 First Midcourse Correction"
  },
  {
    "chunk_id": 148,
    "question": "What was the duration of the planned descent propulsion system maneuver?",
    "answer": "6l-l/2 hours",
    "difficulty": "easy",
    "chunk_content": "A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in \"Auto.\"\n\nPrimary guidance system performance was nomi nal $\\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver."
  },
  {
    "chunk_id": 148,
    "question": "What system was used to maintain attitude after the error needles were nulled?",
    "answer": "primary guidance",
    "difficulty": "medium",
    "chunk_content": "A descent propulsion system maneuver to reestablish a free-return trajectory was planned for 6l-l/2 hours using primary guidance. The docked configuration was maneuvered manually to null out guidance system error needles using the thrust/translation controller assembly for roll and pitch control and the attitude controller assembly for yaw control. It was not difficult to control the docked configuration in this manner. There was, however, some concern as to the effect the use of the thrust/ translation controller assembly would have on the trajectory. After the error needles were nulled, attitude was maintained using primary guidance with attitude control in \"Auto.\"\n\nPrimary guidance system performance was nomi nal $\\cdot$ during the mi dcourse maneuver to a free return. There were no vehicle attitude excursions, and the firing time was as predicted. The abort guidance system was not powered up for this maneuver."
  },
  {
    "chunk_id": 153,
    "question": "What was the initial throttle position?",
    "answer": "idle position",
    "difficulty": "easy",
    "chunk_content": "Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver.\n\n8.9 TRANSEARTH COAST\n\n8.9.1 Coast Phase Activities"
  },
  {
    "chunk_id": 153,
    "question": "What was the final throttle position?",
    "answer": "full throttle",
    "difficulty": "easy",
    "chunk_content": "Guidance system performance was again nominal and there were no significant attitude excursions.. The throttle profile was started in the idle position, then moved to 40 percent for 2l seconds, and finally to full throttle for the remainder of the firing. The maneuver residuals were 0.2, 0.0, and 0.3 ft/sec in the X, Y, and Z axes, respectively. The abort guidance system was powered up and was used to monitor both attitude and velocity change and agreed with primary system readouts throughout the maneuver.\n\n8.9 TRANSEARTH COAST\n\n8.9.1 Coast Phase Activities"
  },
  {
    "chunk_id": 164,
    "question": "How long did it take to fully charge battery B?",
    "answer": "Approximately 3 hours",
    "difficulty": "easy",
    "chunk_content": "The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness."
  },
  {
    "chunk_id": 164,
    "question": "What was the source of power for the command module during the transearth coast?",
    "answer": "The lunar module through the umbilical connectors",
    "difficulty": "medium",
    "chunk_content": "The unprecedented powered-down state of the command module required generation of several new procedures in preparation for entry. The command module was briefly powered up to assess the operation of critical systems using both onboard and telemetered instrumentation. Any required power in the command module had been supplied during transearth coast from the lunar module through the umbilical connectors. It was through this means that the entry batteries were fully charged, with battery A requiring 15 hours and battery B approximately 3 hours. While these procedures represented a radical departure from normal operation, all were understandable and easily accomplished to achieve the desired system readiness."
  },
  {
    "chunk_id": 176,
    "question": "What was the result of the sextant star check and moon occultation?",
    "answer": "Confirmed",
    "difficulty": "easy",
    "chunk_content": "The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g."
  },
  {
    "chunk_id": 176,
    "question": "What was the first acceleration peak during the guided entry?",
    "answer": "Approximately 5g",
    "difficulty": "medium",
    "chunk_content": "The entry attitude and platform alignment were confirmed by a successful sextant star check and moon occulation within l second of the predicted time. The pre-entry check and initialization of the entry monitor system were normal. However, entry monitor system operation was initiated manually when the 0.05g light remained off 3 seconds after the actual $0.05\\mathtt{g}$ time (as discussed in section 14.l.5.). In addition, the entry monitor system trace was unexpectedly narrow and required excessive concentration to read. The guided entry was normal in all respects and was characterized by smooth control inputs. The first acceleration peak reached approximately 5g."
  },
  {
    "chunk_id": 177,
    "question": "What type of flotation equipment is preferable for ease of donning and egress?",
    "answer": "Standard Navy life vests",
    "difficulty": "easy",
    "chunk_content": "Landing decelerations were mild in comparison to Apollo 8, and the spacecraft remained in the stable I flotation attitude after parachute release. Recovery proceeded rapidly and efficiently. Standard Navy life vests were passed to the crew by recovery personnel. For ease of donning and egress, these are preferable to the standard underarm flotation equipment. They would also quite effectively keep an unconscious crewman's head out of the water.\n\n9.0 BIOMEDICAL EVALUATION\n\nThis section is a summary of Apollo l3 medical findings, based on preliminary analyses of biomedical data. From the medical point of view, the first 2 days of the Apollo l3 mission were completely routine. The biomedical data were excellent, and physiological parameters remained within expected ranges. Daily crew status reports indicated that the crewmen were obtaining adequate sleep, no medications were taken, and the radiation dosage was exactly as predicted.\n\n9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA"
  },
  {
    "chunk_id": 177,
    "question": "How long did the first 2 days of the Apollo 13 mission remain routine from a medical point of view?",
    "answer": "Completely",
    "difficulty": "easy",
    "chunk_content": "Landing decelerations were mild in comparison to Apollo 8, and the spacecraft remained in the stable I flotation attitude after parachute release. Recovery proceeded rapidly and efficiently. Standard Navy life vests were passed to the crew by recovery personnel. For ease of donning and egress, these are preferable to the standard underarm flotation equipment. They would also quite effectively keep an unconscious crewman's head out of the water.\n\n9.0 BIOMEDICAL EVALUATION\n\nThis section is a summary of Apollo l3 medical findings, based on preliminary analyses of biomedical data. From the medical point of view, the first 2 days of the Apollo l3 mission were completely routine. The biomedical data were excellent, and physiological parameters remained within expected ranges. Daily crew status reports indicated that the crewmen were obtaining adequate sleep, no medications were taken, and the radiation dosage was exactly as predicted.\n\n9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA"
  },
  {
    "chunk_id": 178,
    "question": "What was the heart rate of the Lunar Module Pilot?",
    "answer": "72",
    "difficulty": "easy",
    "chunk_content": "The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications.\n\nPrior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows.\n\nCrewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12"
  },
  {
    "chunk_id": 178,
    "question": "How many biomedical signals could be monitored in the Lunar Module at a time?",
    "answer": "one",
    "difficulty": "medium",
    "chunk_content": "The biomedical data were excellent in quality during the period from launch to the occurrence of the inflight incident. Physiological data for the remainder of the mission were very scant. The command module was completely powered down, and this eliminated simultaneous biomedical monitoring capability. In the lunar module, only one electrocardiogram signal for one crewman at a time can be monitored. However, even these.medical data were sacrificed to improve air-to-ground communications.\n\nPrior to the abort condition, physiological parameters were well within expected ranges. Just prior to the incident, heart and respiratory rates of the crewmen were as follows.\n\nCrewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12"
  },
  {
    "chunk_id": 179,
    "question": "What was the range of the Command Module Pilot's heart rate during the entry phase?",
    "answer": "60 to 70 beats/min",
    "difficulty": "easy",
    "chunk_content": "At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload.\n\nDuring the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry.\n\n9.2 INFLIGHT HISTORY\n\n9.2.l Adaptation to Weightlessness"
  },
  {
    "chunk_id": 179,
    "question": "What was the indication of the Lunar Module Pilot's heart rate compared to his basal rate?",
    "answer": "an indication of an inflight illness",
    "difficulty": "medium",
    "chunk_content": "At 55:54:54, a telemetry dropout was observed. Immediately after the incident, crew heart rates ranged from $\\mathtt{105}$ to 136 beats /min. These heart rates are well within normal limits and are indicative of stress and an increased workload.\n\nDuring the entry phase, biomedical data on the Command Module Pilot and Lunar Module Pilot were available. The Command Module Pilot's heart rate ranged from 60 to 70 beats/min. The Lunar Module Pilot's heart rate ranged from 100 to l25 beats/min, which in contrast to his basal rate was an indication of an inflight illness detected after flight. The Commander had removed his bioharness shortly after the emergency incident; hence, no biomedical data were available from him during the entry.\n\n9.2 INFLIGHT HISTORY\n\n9.2.l Adaptation to Weightlessness"
  },
  {
    "chunk_id": 184,
    "question": "What was the nickel content in the command module hot water port?",
    "answer": "$0.05~\\mathrm{mg/1}$",
    "difficulty": "easy",
    "chunk_content": "Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again."
  },
  {
    "chunk_id": 184,
    "question": "What was the outcome of the bacterial population after inflight chlorination?",
    "answer": "to specification levels",
    "difficulty": "medium",
    "chunk_content": "Preflight testing of both command module and lunar module water supplies revealed no significant contaminants. The nickel content from samples taken at the command module hot water port was $0.05~\\mathrm{mg/1}$ .Elevated nickel concentration has been a consistent finding in previous missions and has been ruled acceptable in view of no detrimental effects on crew physiology. There was a substantial buildup in total bacterial count from the time of final filling of the command module potable water system until final preflight sampling 24 hours prior to launch. This count was deemed acceptable under the assumption the first inflight chlorination would reduce the bacterial population to specification levels. Preflight procedures will be reviewed to investigate methods of preventing growth of organisms in the command module water system during the countdown phase. The inflight chlorination schedule was followed prior to the incident, after which the potable water was not chlorinated again."
  },
  {
    "chunk_id": 185,
    "question": "How many ounces of water were consumed by each crewman after the oxygen tank incident?",
    "answer": "24",
    "difficulty": "easy",
    "chunk_content": "The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress.\n\n9.2.5 Food\n\nThe flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches."
  },
  {
    "chunk_id": 185,
    "question": "What was the basis for selecting the flight menus?",
    "answer": "Crew preferences determined by preflight evaluation of representative flight foods",
    "difficulty": "medium",
    "chunk_content": "The crew rationed water and used it sparingly after the oxygen tank incident. Not more than 24 ounces of water were consumed by each crewman after the incident. The crew reported that the juice bags contained about 20 percent gas, but that this amount was not enough to cause any distress.\n\n9.2.5 Food\n\nThe flight menus were similar to those of prior Apollo missions and were designed to provide approximately 2lo0 kilocalories per man per day. The menus were selected on the basis of crew preferences determined by preflight evaluation of representative flight foods. There were no modifications to the menu as a result of the late crew change. New food items for this mission included meatballs with sauce, cranberry-orange relish, chicken and rice soup, pecans, natural orange juice crystals, peanut butter, and jelly. Mustard and tomato catsup were also provided for the sandwiches."
  },
  {
    "chunk_id": 191,
    "question": "What were the descriptions given by the crewmen for the flashes of light they observed?",
    "answer": "\"pinpoint novas ,\" \"roman candles,\" and \"similar to traces in a cloud chamber.\"",
    "difficulty": "easy",
    "chunk_content": "The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as \"pinpoint novas ,\" \"roman candles,\" and \"similar to traces in a cloud chamber.\" More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute.\n\n9.3 PHYSICAL EXAMINATIONS\n\nPreflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary."
  },
  {
    "chunk_id": 191,
    "question": "How many physical examinations were conducted for the primary crew before launch?",
    "answer": "3",
    "difficulty": "easy",
    "chunk_content": "The crew reported seeing point flashes or streaks of light, as had been previously observed by the Apollo ll and l2 crews. The crewmen were aware of these flashes only when relaxed, in the dark, and with their eyes closed. They described the flashes as \"pinpoint novas ,\" \"roman candles,\" and \"similar to traces in a cloud chamber.\" More point flashes than streaks were observed, and the color was always white.. Estimates of the frequency ranged from 4 flashes per hour to 2 flashes per minute.\n\n9.3 PHYSICAL EXAMINATIONS\n\nPreflight physical examinations of both the primary and backup crews were conducted 30 days prior to launch, and examinations of the primary crew only were conducted 15 and 5 days prior to laumch. The Lunar Module Pilot suffered a sore throat 18 days before launch, and throat swabs from all three crewmen were cultured on two occasions. Since the organism identified was not considered pathogenic and the crew showed no symptoms of illness , no treatment was necessary."
  },
  {
    "chunk_id": 192,
    "question": "How many days before flight did the primary Command Module Pilot get exposed to rubella?",
    "answer": "Eight",
    "difficulty": "easy",
    "chunk_content": "Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations."
  },
  {
    "chunk_id": 192,
    "question": "What was the result of the laboratory studies on the primary Command Module Pilot?",
    "answer": "No immunity to rubella",
    "difficulty": "medium",
    "chunk_content": "Eight days before flight, the primary Command Module Pilot was exposed to rubella (German measles) by a member of the backup crew. The physical examination 5 days before flight was normal, but laboratory studies revealed that the primary Command Module Pilot had no immunity to rubella. Consequently, on the day prior to launch the final decision was made to replace the primary Command Module Pilot with the backup Command Module Pilot. A complete physical examination had been conducted on the backup Command Module Pilot 3 days before flight, and no significant findings were present in any preflight histories or examinations."
  },
  {
    "chunk_id": 193,
    "question": "What was the weight loss of the Lunar Module Pilot?",
    "answer": "6.5 pounds",
    "difficulty": "easy",
    "chunk_content": "Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses.\n\nThe Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites.\n\n10.0 MISSION SUPPORT PERFORMANCE\n\n10.1 FLIGHT CONTROL"
  },
  {
    "chunk_id": 193,
    "question": "What was the cause of the Lunar Module Pilot's dizziness?",
    "answer": "Fatigue, the effects of weightlessness, and the urinary tract infection",
    "difficulty": "medium",
    "chunk_content": "Postflight physical examinations were conducted immediately after recovery. These physical examinations were normal, although all crewmen were extremely fatigued and the Lunar Module Pilot had a urinary tract infection. While standing during portions of his postflight physical examination, the Lunar Module Pilot had several episodes of dizziness, which were attributed to fatigue, the effects of weightlessness, and the urinary tract infection. The Commander, Command Module Pilot, and Lunar Module Pilot exhibited weight losses of 14, ll, and 6.5 pounds , respectively. In the final 4 or 5 hours of the flight, the Lunar Module Pilot drank considerably more water than did the other crewmen and possibly replenished his earlier body fluid losses.\n\nThe Command Module Pilot had a slight irritation at the site of the superior sensor on the upper chest, but the Cormander and Lunar Module Pilot had no irritation at any sensor sites.\n\n10.0 MISSION SUPPORT PERFORMANCE\n\n10.1 FLIGHT CONTROL"
  },
  {
    "chunk_id": 195,
    "question": "What was the primary mission decision after the tanks containing cryogenic oxygen experienced a problem?",
    "answer": "Abort the primary mission and attempt a safe return to earth as rapidly as possible.",
    "difficulty": "easy",
    "chunk_content": "The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions:\n\na. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers"
  },
  {
    "chunk_id": 195,
    "question": "What system was used for life support after the oxygen tanks experienced a problem?",
    "answer": "Lunar module",
    "difficulty": "medium",
    "chunk_content": "The two tanks cortaining cryogenic oxygen, used for _fuel cell operation and crew breathing, experienced a problem at about 56 hours, as described in section l4.l.l and reference l. This condition resulted in the following flight control decisions:\n\na. Abort the primary mission and attempt a safe return to earth as rapidly as possible. b. Shut down all command and service module systems to conserve consumables for entry. c. Use the lunar module for life support and any propulsive maneuVers"
  },
  {
    "chunk_id": 196,
    "question": "What time was the power down of the command and service nodules and power up of the lunar module completed?",
    "answer": "58:40:00",
    "difficulty": "easy",
    "chunk_content": "Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always"
  },
  {
    "chunk_id": 196,
    "question": "What type of maneuver was used to expedite the landing to about 142:30:00?",
    "answer": "pericynthion-plus-2-hour maneuver (transearth injection)",
    "difficulty": "medium",
    "chunk_content": "Powering down of the command and service nodules and powering up of the lunar module were completed at 58:40:00. The optimum plan for 8 safe and quick return required an immediate descent engine firing to a free-return circumlunar trajectory, with a pericynthion-plus-2-hour maneuver (transearth injection) to expedite the landing to about 142:30:00. Two other midcourse corrections were performed, the first using the descent engine. Only essential life support, navigation, instrumentation, and communication systems were operated to maximize electrical power and cooling water margins. Detailed monitoring of all consumables was continuously maintained to assess these margins, and the crew was always"
  },
  {
    "chunk_id": 197,
    "question": "What system was used by the crew to remove carbon dioxide in the lunar module?",
    "answer": "Environmental control system",
    "difficulty": "easy",
    "chunk_content": "advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast."
  },
  {
    "chunk_id": 197,
    "question": "What was established using the lunar module reaction control system?",
    "answer": "Passive thermal control mode",
    "difficulty": "medium",
    "chunk_content": "advised of their consumables status. A procedure was developed on the ground and used by the crew to allow use of command module lithium hydroxide cartridges for carbon dioxide removal in the lunar module environmental control system (see section 6.8). The passive thermal control mode was established using the lunar module reaction control system and was satisfactorily maintained throughout transearth coast."
  },
  {
    "chunk_id": 198,
    "question": "What was the task that was verified in a simulator before advising the crew?",
    "answer": "procedures",
    "difficulty": "easy",
    "chunk_content": "A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target.\n\n10.2 NETWORK"
  },
  {
    "chunk_id": 198,
    "question": "How long before entry did lunar module undocking occur?",
    "answer": "1 hour",
    "difficulty": "medium",
    "chunk_content": "A major flight control function, in adaition to the monitoring of systems status and maintaining of consumable quantities above red-line values, was to determine the procedures to be used immediately prior to and during entry. After satisfactory procedures were established, they were verified in a simulator prior to advising the crew. These procedures called for first separating the service module, remaining on lunar module environmental control and power as late as possible, coaligning the two platforms, and separating the lunar module using tunnel pressure. The command module tunnel hatch was installed and a leak check was performed prior to lunar module undocking, which occurred about 1 hour before entry. All spacecraft operations were normal from undocking through landing, which occurred very close to the established target.\n\n10.2 NETWORK"
  },
  {
    "chunk_id": 203,
    "question": "What type of aircraft is capable of lifting the command module?",
    "answer": "HH-53C Helicopters",
    "difficulty": "easy",
    "chunk_content": "TABLE 1O.3-I.- RECOVERY SUPPORT\n\nLanding area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, \u2019Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension  Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii\n\n&arotal ship support $=5$ Total aircraft support $\\approx$ 23\n\n10.3.l Command Module Location and Retrieval"
  },
  {
    "chunk_id": 203,
    "question": "How many SH-3 Helicopters are staged from Norfolk NAS, Virginia?",
    "answer": "1",
    "difficulty": "easy",
    "chunk_content": "TABLE 1O.3-I.- RECOVERY SUPPORT\n\nLanding area Supporta Remarks Number Unit Launch site 1 LCU Landing craft utility (landing craft with command Launch abort 1 HH-3E module retrieval capability)- USS Paiute Helicopter with para-rescue team staged from Patrick AFB,Florida 2 HH-53C Helicopters capable of lifting the command module; each with para-rescue team staged from Patrick AFB, 1 ATF Florida 2 SH-3 Helicopters staged from Norfolk NAS, Virginia 1 DD USS New 3 HC-130H Fixed wing aircraft; one each staged from McCoy AFB, Florida; Pease AFB,New Mexico; and Lajes AFB, \u2019Azores Earth orbit Primary end-of-mission, 2 2 DD HC-130H USS New Fixed wing aircraft staged from Ascension  Mid-Pacific earth 1 1 HdT DD USS Iwo Jima 8 SH-3D USS Benjamin Stoddert orbital, and deep- 2 Helicopters staged from USS Iwo Jima space secondary HC-130H Fixed wing aircraft staged from Hickam AFB, Hawaii\n\n&arotal ship support $=5$ Total aircraft support $\\approx$ 23\n\n10.3.l Command Module Location and Retrieval"
  },
  {
    "chunk_id": 208,
    "question": "Where was the flight crew flown to after being flown from Pago Pago, Samoa?",
    "answer": "Hawaii",
    "difficulty": "easy",
    "chunk_content": "The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston.\n\nUpon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft.\n\nThe following is a chronological listing of events during the recovery operations."
  },
  {
    "chunk_id": 208,
    "question": "How many days did it take for the command module to be flown from Hickam Air Force Base to the manufacturer's plant at Downey, California?",
    "answer": "Two and one half",
    "difficulty": "medium",
    "chunk_content": "The flight crew remained aboard the Iwo Jima overnight and were flown to Pago Pago, Samoa, the following morning. A C-l4l aircraft then took the crew to Hawaii, and following a ceremony and an overnight stay, they were returned to Houston.\n\nUpon arrival of the Iwo Jima in Hawaii, the command module was offloaded and taken to Hickam Air Force Base for deactivation. Two and one half days later, the command module was flown to the manufacturer's plant at Downey, California aboard a C-l33 aircraft.\n\nThe following is a chronological listing of events during the recovery operations."
  },
  {
    "chunk_id": 209,
    "question": "When was the flotation collar inflated?",
    "answer": "1824",
    "difficulty": "easy",
    "chunk_content": "Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928\n\n10.3.2 Postrecovery Inspection"
  },
  {
    "chunk_id": 209,
    "question": "What was completed at 1936 on April 18?",
    "answer": "Flight crew departed Iwo Jima",
    "difficulty": "easy",
    "chunk_content": "Event Time, G.m.t. Apri1 17, 1970 S-band contact by Samoa Rescue 4 Visual contact by Swim 2 1801 1802 helicopters Voice contact by Recovery helicopter 1803 Visual contact by Relay/Recovery helicopters/ 1803 Iwo Jima Command module landed, remained in stable I Swimmers deployed to retrieve main parachutes 1807 1809 First swimmer deployed to command module 1816 Flotation collar inflated 1824 Life preserver unit delivered to lead swimmer 1831 Command module hatch opened 1832 Helicopter pickup of flight crew completed 1842 Recovery helicopter on board Iwo Jima 1853 Command module secured aboard Iwo Jima 1936 April 18 Flight crew departed Iwo Jima 1820 April 20 Flight crew arrival in Houston 0330 Iwo Jima arrival in Hawaii April 24 1930 Safing of command module pyrotechnics completed April_25 0235 Deactivation of the fuel and oxidizer completed April 26 1928\n\n10.3.2 Postrecovery Inspection"
  },
  {
    "chunk_id": 211,
    "question": "What was the condition of the interior surfaces of the command module?",
    "answer": "very damp and cold",
    "difficulty": "easy",
    "chunk_content": "a. Some of the radioluminescent disks were broken. b. The apex cover was broken on the extravehicular handle side. c. The docking ring was burned and broken. d. The right--hand roll thruster was blistered. e. A yellowish/tan film existed on the outside of the hatch window, left and right rendezvous windows, and the right-hand window. f. The interior surfaces of the command module were very damp and cold, assumed to be condensation; there was no pooling of water on the floor. . Water samples could not be taken from the spacecraft tanks (discussed in section 5.8). h. The postlanding ventilation exhaust valve was open and the inlet valve was closed; the postlanding ventilation valve unlock handle was apparently jammed between the lock and unlock positions (section 14.l.2). i. There was more and deeper heat streaking in the area of the compression and shear pads than has been normally observed.\n\n11.0 EXPERIMENTS\n\n11.1 ATMOSPHERIC ELECTRICAL PHENOMENA"
  },
  {
    "chunk_id": 211,
    "question": "What was the state of the postlanding ventilation valve unlock handle?",
    "answer": "jammed between the lock and unlock positions",
    "difficulty": "medium",
    "chunk_content": "a. Some of the radioluminescent disks were broken. b. The apex cover was broken on the extravehicular handle side. c. The docking ring was burned and broken. d. The right--hand roll thruster was blistered. e. A yellowish/tan film existed on the outside of the hatch window, left and right rendezvous windows, and the right-hand window. f. The interior surfaces of the command module were very damp and cold, assumed to be condensation; there was no pooling of water on the floor. . Water samples could not be taken from the spacecraft tanks (discussed in section 5.8). h. The postlanding ventilation exhaust valve was open and the inlet valve was closed; the postlanding ventilation valve unlock handle was apparently jammed between the lock and unlock positions (section 14.l.2). i. There was more and deeper heat streaking in the area of the compression and shear pads than has been normally observed.\n\n11.0 EXPERIMENTS\n\n11.1 ATMOSPHERIC ELECTRICAL PHENOMENA"
  },
  {
    "chunk_id": 213,
    "question": "How many electric field meters were installed in the area to the north and west of the launch site?",
    "answer": "Nine",
    "difficulty": "easy",
    "chunk_content": "As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges."
  },
  {
    "chunk_id": 213,
    "question": "What was the purpose of the special device operated at site 5?",
    "answer": "To measure rapid changes in the electric field",
    "difficulty": "medium",
    "chunk_content": "As shown in figures ll.l-l and ll.l-2, a network of nine calibrated electric field meters was installed in the area to the north and west of the launch site. Seven of the field meters were connected to multiple channel recorders so that any excursions of the electric field intensity could be measured over a wide range of yalues. A special device was operated at site 5, located on the beach 4 miles northwest of the launch site. This device was installed to measure rapid changes in the electric field and was used, together with a sferics detector, to sense the electromagnetic radiation generated by lightning or other significant electrical discharges."
  },
  {
    "chunk_id": 215,
    "question": "What was the nature of the initial change in the record at site 6?",
    "answer": "Positive",
    "difficulty": "easy",
    "chunk_content": "At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by\n\n\n\nFigure ll.l-l.- Field meter location in the laumch site area.\n\n\n\nFigure ll.l-2.-- Field meter locations in the proximity of the launch complex.\n\ngravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6."
  },
  {
    "chunk_id": 215,
    "question": "What was the approximate negative field value at site 6 after 40 seconds?",
    "answer": "3000 volts/meter",
    "difficulty": "medium",
    "chunk_content": "At site 6, the record was similar to that for site 7 with an initial positive excursion followed by a slower negative change. At this station, however, there were large fluctuations superimposed on the record, as shown in figure ll.l-3(b). These fluctuations could have been caused by\n\n\n\nFigure ll.l-l.- Field meter location in the laumch site area.\n\n\n\nFigure ll.l-2.-- Field meter locations in the proximity of the launch complex.\n\ngravel and dust stirred up by the exhaust of the launch vehicle engine. After launch, a quantity of such debris was found near the surface of the field meter and its surrounding area. After the oscillations had subsided at T plus 40 seconds, there was a large negative field of approximately minus 3000 volts/meter which probably resulted from the exhaust and steam clouds that tended to remain over site 6."
  },
  {
    "chunk_id": 216,
    "question": "What was the reason for the recorders at sites 8 and 9 being started several hours prior to launch?",
    "answer": "Access restrictions to sites 8 and 9",
    "difficulty": "medium",
    "chunk_content": "Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off.\n\n\n\nFigure ll.l-3.- Electrical discharge data for the Apollo 13 launch.\n\n\n\nFigure 1l.l-3.- Concluded\n\nNo significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations."
  },
  {
    "chunk_id": 216,
    "question": "What was the source of the large field perturbations found on the stationary parts of the records at sites 8 and 9?",
    "answer": "The launch",
    "difficulty": "easy",
    "chunk_content": "Because of access restrictions to sites 8 and 9, the corresponding recorders were started several hours prior to launch and unfortumately had stopped before lift-off. However, substantial positive and negative field perturbations found on the stationary parts of the records were greater than anything found on the moving portion. Comparison of these records with those from sites 6 and 7 confirmed that the only large field perturbations were those accompanying launch. Consequently, the peak excursions of the records at sites 8 and 9 could be confidently associated with the maximum field perturbations occurring just after lift-off.\n\n\n\nFigure ll.l-3.- Electrical discharge data for the Apollo 13 launch.\n\n\n\nFigure 1l.l-3.- Concluded\n\nNo significant perturbation in the electric field was produced by the launch cloud at stations 4 or 5, although small-scale fluctuations, apparently resulting from vibrations, can be seen on the records of the fine weather field at both stations."
  },
  {
    "chunk_id": 217,
    "question": "Did the field-change and sferics detectors at site 5 detect any lightning-like discharge during the launch?",
    "answer": "No",
    "difficulty": "easy",
    "chunk_content": "The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12."
  },
  {
    "chunk_id": 217,
    "question": "What was the origin of the sporadic signals recorded during the afternoon of launch day?",
    "answer": "Lightning in a cold front",
    "difficulty": "medium",
    "chunk_content": "The field-change and sferics detectors at site 5 gave no indication of any lightning-like discharge during launch, although sporadic signals were later recorded during the afternoon of launch day. These signals probably came from lightning in a cold front which was stalled some distance to the northwest of the launch site and which passed over the launch site on April 12."
  },
  {
    "chunk_id": 219,
    "question": "What is the equilibrium potential for a conventional jet aircraft?",
    "answer": "a million volts",
    "difficulty": "easy",
    "chunk_content": "The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V.\n\nll.l.2 Very-Low and Low-Frequency Radio Noise"
  },
  {
    "chunk_id": 219,
    "question": "What is the estimated electrostatic potential of a Saturn V?",
    "answer": "several million volts",
    "difficulty": "medium",
    "chunk_content": "The electrostatic potential developed on an aircraft is caused by the engine charging current, which, in turn, is balanced by the corona current loss from the aircraft. For a conventional jet aircraft, this equilibrium potential approaches a million volts. For the Saturn V launch vehicle, the charging current probably is far greater than that of a jet aircraft. Furthermore, since the surface of an aircraft probably has more external irregularities than a launch vehicle, the charging current is higher and the corona current loss is typically less for a launch vehicle than for an aircraft. Both of these effects tend to make the equilibrium potential for the Saturn vehicle larger than that of 8 jet aircraft; therefore, several million volts does not seem to be an unreasonable estimate for the electrostatic potential of a Saturn V.\n\nll.l.2 Very-Low and Low-Frequency Radio Noise"
  },
  {
    "chunk_id": 220,
    "question": "What was the frequency to which the first receiver was tuned?",
    "answer": "1.5 kHz",
    "difficulty": "easy",
    "chunk_content": "To monitor the low-frequency radio noise, a broad-band antenna system was used at site 7 to feed five receivers, tuned respectively to 1.5 kHz, 6 kHz, 27 kHz, 51 kHz, and 120 kHz.\n\nDuring launch, a sudden onset of radio noise was observed almost coincidently with the start of the electric field perturbation. This onset was very well marked on all but the 1.5 kHz channel. Following onset, the noise levels at l20 and at 5l kHz tended to decrease slowly in intensity for some 20 seconds. However, the noise levels at 27 and at 6 kHz increased and reached their maxima after about 15 seconds. Furthermore, substantial noise at 1.5 kHz was first apparent at 5 seconds after lift-off and also peaked out in about l5 secorids."
  },
  {
    "chunk_id": 220,
    "question": "How long did the noise levels at 27 and 6 kHz take to reach their maxima after the onset of noise?",
    "answer": "15 seconds",
    "difficulty": "medium",
    "chunk_content": "To monitor the low-frequency radio noise, a broad-band antenna system was used at site 7 to feed five receivers, tuned respectively to 1.5 kHz, 6 kHz, 27 kHz, 51 kHz, and 120 kHz.\n\nDuring launch, a sudden onset of radio noise was observed almost coincidently with the start of the electric field perturbation. This onset was very well marked on all but the 1.5 kHz channel. Following onset, the noise levels at l20 and at 5l kHz tended to decrease slowly in intensity for some 20 seconds. However, the noise levels at 27 and at 6 kHz increased and reached their maxima after about 15 seconds. Furthermore, substantial noise at 1.5 kHz was first apparent at 5 seconds after lift-off and also peaked out in about l5 secorids."
  },
  {
    "chunk_id": 223,
    "question": "At what time were the first two balloons launched?",
    "answer": "April 9, 1970",
    "difficulty": "easy",
    "chunk_content": "Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the \"fair weather\" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an\n\naltitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood.\n\n11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES"
  },
  {
    "chunk_id": 223,
    "question": "What was the frequency of the oscillating current before the marked increase in amplitude?",
    "answer": "15 cycles per minute",
    "difficulty": "medium",
    "chunk_content": "Three balloons containing instruments designed to measure the air/ earth current density were launched: at 6:52 p.m. on April 9, 1970, and at 1:14 p.m. and 1:52 p.m. on April 1l, l970. The first two ba1loons provided the \"fair weather\" base for the experiment. At lift-off, the third balloon was about l2.2 miles southeast of the launch site at an\n\naltitude of 20 0o0 feet. Forty-five seconds after lift-off, the current density, which had been oscillating at a frequency of about 15 cycles per minute, showed a marked increase in amplitude. This variation in current was again observed when the balloon reached an altitude between 40 000 and 50 000 feet. The frequency of the observed current variation was also noted from the balloon released at l:l4 p.m. The cause of the oscillating current and the enhancement thereof are not yet understood.\n\n11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES"
  },
  {
    "chunk_id": 226,
    "question": "How many photographs were taken for the test?",
    "answer": "Eleven",
    "difficulty": "easy",
    "chunk_content": "To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately,"
  },
  {
    "chunk_id": 226,
    "question": "What was required to reconstruct the geometry involved in the test?",
    "answer": "A precise record of time of photography",
    "difficulty": "medium",
    "chunk_content": "To aid in a test of the feasibility of performing stereoscopic determination of cloud height at synchronous altitudes, a series of earthcentered photographs at 20-minute intervals, beginning soon after translunar injection, were planned. The photographs required for this test could only have been acquired from an Apollo lunar mission. A precise record of time of photography was required to reconstruct the geometry involved. Eleven photographs were taken, and a precise time record was obtained. The description of the location of the spacecraft at the time of each photograph is given in table ll.2-I, along with the time of photography, the enlargement required on each frame for normalization, and the distance between photographic points. The experiment was successful, and all photographs are of excellent quality. To support the analysis of these photographs, aircraft reports, synoptic weather charts and satellite photographs for the time of photography have been acquired. Unfortunately,"
  },
  {
    "chunk_id": 228,
    "question": "What is the altitude of the object in the 13-60-8594 frame?",
    "answer": "1982.8",
    "difficulty": "easy",
    "chunk_content": "TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY\n\nMagazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130\u00b000'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28\u00b025'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27\u00b039'N 151\u00b039*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156\u00b035'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27\u00b014'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27\u00b004+N 165\u00b09*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26\u00b054'N 170\u00b050'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175\u00b051'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26\u00b036'N 179\u00b014*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26\u00b027'N g60 56728 9.251 1.6254 2436.6"
  },
  {
    "chunk_id": 228,
    "question": "What is the latitude of the object in the 13-60-8591 frame?",
    "answer": "28\u00b025'N",
    "difficulty": "easy",
    "chunk_content": "TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY\n\nMagazine L frame Mission elapsed time hr:min:sec Gmt hr:min:sec Latitude Longitude Altitude Normalization enlargement required Distance apart mile Mile Earth radii (from center) 13-60-8590 07:17:14 02:30:46 28038/Na 130\u00b000'wa 006 6.076 1.00000 13-60-8591 07:39:47 02:52:49 28\u00b025'N 37054 6.389 1.0617 1473.5 13-60-8592 08:42:07 03:55:09 270491Na 147030'wa 180 7.280 1.2372 4409.2 13-60-8593 09:03:11 04:16:13 27\u00b039'N 151\u00b039*W 44 998 7.545 1.2893 1609.5 13-60-8594 09:26:34 04:29:36 156\u00b035'W 47 098 7.850 1.3495 1982.8 13-60-8595 09:47:10 05:00:12 27\u00b014'Na 161000 48 920 8.116 1.4017 1848.0 13-60-8596 10:08:39 05:21:41 27\u00b004+N 165\u00b09*W 49 876 8.255 1.4291 2240.4 13-60-8597 10 :30:59 05:44:01 26\u00b054'N 170\u00b050'W 51 655 8.513 1.4800 2202.6 13-60-8598 10 : 52 : 59 06:06:01 260451a 175\u00b051'W 53 TOt 8.767 1.5301 2275.5 13-60-8599 11:14:59 06:28:01 26\u00b036'N 179\u00b014*E 55 056 9.008 1.5775 2296.8 13-60-8600 11 : 37 : 19 06:50:21 26\u00b027'N g60 56728 9.251 1.6254 2436.6"
  },
  {
    "chunk_id": 231,
    "question": "How long did seismic signals continue after impact?",
    "answer": "over 4 hours",
    "difficulty": "easy",
    "chunk_content": "Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds."
  },
  {
    "chunk_id": 231,
    "question": "What was the peak signal intensity compared to the Apollo 12 ascent stage impact?",
    "answer": "8 times larger",
    "difficulty": "medium",
    "chunk_content": "Seismic signals were first recorded 28.4 seconds after impact and continued for over 4 hours. Some sigmals were so large that seismometer sensitivity had to be reduced by command from earth to keep the data on Scale. Peak signal intensity occurred l0 minutes after initial onset. The peak value was 8 times larger than that recorded from the Apollo 12 ascent stage impact, which occurred at a range of 40 miles from the seismic station and was equivalent to l ton of rNT. An expanding gas cloud, which presumably swept out over the lunar surface from the S-IVB impact point, was recorded by the lunar ionosphere detector deployed during Apollo 12. Detection of this cloud began approximately 8 seconds before the first seismic signal and lasted 70 seconds."
  },
  {
    "chunk_id": 232,
    "question": "What is the estimated depth range of the S-IVB seismic energy in the moon?",
    "answer": "20 to 40 kilometers",
    "difficulty": "easy",
    "chunk_content": "The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon."
  },
  {
    "chunk_id": 232,
    "question": "What is the velocity at which the initial signal from the S-IVB impact travelled to the seismic station?",
    "answer": "4.8 km/sec",
    "difficulty": "easy",
    "chunk_content": "The character of the signal from the S-IVB impact is identical to that of the ascent stage impact and those from natural events, presumed to be meteoroid impacts, which are being recorded at the rate of about one per day. The S-IVB seismic energy is believed to have penetrated into the moon to a depth of from 20 to 40 kilometers. The initial signal was unusually clear and travelled to the seismic station at a velocity of 4.8 km/sec, which is near that predicted from laboratory measurements using Apollo l2 lunar rock samples. This result implies that, to depths of at least 20 kilometers, the moon's outer shell may be formed from the same crystalline rock material as found at the surface. No evidence of a lower boundary to this material has been found in the seismic signal, although it is clear the material is too dense to form the entire moon."
  },
  {
    "chunk_id": 234,
    "question": "What was the purpose of deploying and activating an Apollo lunar surface experiments package?",
    "answer": "c. Further develop man's capability to work in the lunar environment.",
    "difficulty": "easy",
    "chunk_content": "a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites.\n\nThirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned:\n\nTABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS"
  },
  {
    "chunk_id": 234,
    "question": "How many detailed objectives were derived from the four primary objectives?",
    "answer": "Thirteen",
    "difficulty": "medium",
    "chunk_content": "a. Perform selenological inspection, survey, and sampling of materials in a preselected region of the Fra Mauro formation. b. Deploy and activate an Apollo lunar surface experiments package. c. Further develop man's capability to work in the lunar environment. d. Obtain photographs of candidate exploration sites.\n\nThirteen detailed objectives, listed in table l2-I and described in reference 8, were derived from the four primary objectives. None of these objectives were accomplished because the mission was aborted. In addition to the spacecraft and lunar surface objectives, the following two launch vehicle secondary objectives were assigned:\n\nTABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS"
  },
  {
    "chunk_id": 235,
    "question": "What was the purpose of the S-059 experiment?",
    "answer": "Lunar field geology",
    "difficulty": "easy",
    "chunk_content": "Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes"
  },
  {
    "chunk_id": 235,
    "question": "Was the pilot involved in describing a function during the mission?",
    "answer": "Yes",
    "difficulty": "easy",
    "chunk_content": "Description Completed B C Television coverage No Contingency sample collection Selected sample collection No No D Evaluation of landing accuracy techniques No F No G Photographs of candidate exploration sites H Extravehicular communication performance No No I Lunar soil mechnics No J Dim light photography K Selenodetic reference point update No CSM orbital. science photography No L Transearth lunar photography No M EMU water consumption measurement No N Thermal coating degradation No ALSEPIII Apollo lunar surface experiments package No S-059 Lunar field geology No S-080 Solar wind composition No S-164 S-band transponder exercise No S-170 Downlink bistatic radar observations of the Moon No S-178 Gegenschein from lunar orbit No S-184 Lunar surface close-up photography No T-029 Pilot describing function Yes"
  },
  {
    "chunk_id": 238,
    "question": "What was the initial azimuth of the vehicle launch?",
    "answer": "90 degrees east of north",
    "difficulty": "easy",
    "chunk_content": "The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target"
  },
  {
    "chunk_id": 238,
    "question": "How many seconds later than predicted did the event of orbital insertion occur?",
    "answer": "44.07",
    "difficulty": "medium",
    "chunk_content": "The vehicle was launched on an azimuth 90 degrees east of north, and a roll maneuver at l2.6 seconds placed the vehicle on a flight azimuthof $72.043$ degrees east of north. Trajectory parameters were close to nominal during S-IC and S-II boost until early shutdown of the center engine. The premature cutoff caused considerable deviations from certain nominal launch-vehicle trajectory parameters which were particularly evident at S-II outboard engine cutoff. Despite these deviations, the guidance system is designed to operate such that an efficient boost is conducted under engine-out conditions, and near-nominal trajectory parameters were achieved at orbital insertion and at translunar injection. Because of the reduced effective thrust, however, these respective events occurred $44.07$ and 13.56 seconds later than predicted. After spacecraft ejection, various S-IVB attitude and propulsive maneuvers placed the vehicle on a lunar impact trajectory very close to the desired target"
  },
  {
    "chunk_id": 240,
    "question": "What was the percentage of the maximum lateral loads experienced during S-IC boost compared to the design value?",
    "answer": "25 percent",
    "difficulty": "easy",
    "chunk_content": "Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of \u00b1225 psi initiated engine cutoff through the \"thrust OK\" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an"
  },
  {
    "chunk_id": 240,
    "question": "At what time did the S-II crossbeam oscillations reach a peak amplitude of +33.7g?",
    "answer": "330.6 seconds",
    "difficulty": "medium",
    "chunk_content": "Structural loads experienced during S-IC boost were well below design values, with maximum lateral loads approximately 25 percent of the design value. As a result of high amplitude longitudinal oscillations during S-Il boost, the center engine experienced a l32-second premature cutoff. At 330.6 seconds, the S-II crossbeam oscillaticns reached a peak amplitude of +33.7g: Corresponding center-engine chamber pressure oscillations of \u00b1225 psi initiated engine cutoff through the \"thrust OK\" switches. These responses were the highest measured amplitude for any S-Il flight. Except for the unexpected high amplitude, oscillations in this range are an inherent characteristic of the present S-II structure/propulsion configuration and have been experienced on previous flights. Acceleration levels experienced at various vehicle stations during the period of peak oscillations indicate that the vehicle did not transmit the large magnitude oscillations to the spacecraft. Installation of an"
  },
  {
    "chunk_id": 242,
    "question": "What was the time when the crew heard and felt the vibrations from a sharp \"bang\" during the Apollo 13 mission?",
    "answer": "55 hours 55 minutes",
    "difficulty": "easy",
    "chunk_content": "The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted.\n\n14.0 ANOMALY SUMMARY\n\nThis section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission.\n\n14.1 COMMAND AND SERVICE MODULES\n\n14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure\n\nAt approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp \"bang,\" coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system:"
  },
  {
    "chunk_id": 242,
    "question": "What was the condition associated with the master alarm that occurred during the Apollo 13 mission?",
    "answer": "Main-bus-B undervoltage condition",
    "difficulty": "medium",
    "chunk_content": "The pilot describing function experiment (T-029) was a success, in that data were obtained during manually controlled spacecraft maneuvers which are available to the principle investigator. None of the other experiments was attempted.\n\n14.0 ANOMALY SUMMARY\n\nThis section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission.\n\n14.1 COMMAND AND SERVICE MODULES\n\n14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure\n\nAt approximately 55 hours 55 minutes into the Apollo 13 mission, the crew heard and felt the vibrations from a sharp \"bang,\" coincident with a computer restart and a master alarm associated with a main-bus-B undervoltage condition. Within 20 seconds, the crew made an immediate verification of electrical-system parameters, which appeared normal. However, the crew reported the following barberpole indications from the Service module reaction control system:"
  },
  {
    "chunk_id": 244,
    "question": "What was the pressure at which the tank-2 relief-valve full-flow conditions occurred?",
    "answer": "1008 psia",
    "difficulty": "easy",
    "chunk_content": "Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers."
  },
  {
    "chunk_id": 244,
    "question": "How long did the pressure decrease last before the relief valve probably reseated?",
    "answer": "9 seconds",
    "difficulty": "medium",
    "chunk_content": "Electrical shorts in the fan circuit ignited the wire insulation, causing pressure and temperature increases within oxygen tank 2. During the pressure rise period, the fuses opened in both fan circuits in cryogenic oxygen tank 2. A short-circuit conduction in the quantity gaging system cleared itself and then began an open-circuit condition. When the pressure reached the tank-2 relief-valve full-flow conditions of 1008 psia, the pressure decreased for about 9 seconds, after which time the relief valve probably reseated, causing another momentary pressure increase. Approximately $1/4$ second after this momentary pressure increase, a vibration disturbance was noted on the command module accelerometers."
  },
  {
    "chunk_id": 245,
    "question": "What was the immediate cause of the tank line to burst?",
    "answer": "Burning of the wire insulation",
    "difficulty": "easy",
    "chunk_content": "The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as"
  },
  {
    "chunk_id": 245,
    "question": "How long was data lost due to the panel separation shock?",
    "answer": "1.8 seconds",
    "difficulty": "easy",
    "chunk_content": "The next series of events occurred within a fraction of a second between the accelerometer disturbances and a momentary loss of data. Burning of the wire insulation reached the electrical conduit leading from inside the tube to the external plug causing the tank line to burst because of overheating. The ruptured electrical conduit caused the vacuum jacket to over pressurize and, in turn, caused the blow-out plug in the vacuum jacket to rupture. Some mechanism, possibly the burning of insulation in bay 4 combined with the oxygen buildup in that bay, caused a rapid pressure rise which resulted in separation of the outer panel. Ground tests, however, have not substantiated the burning of the Mylar insulation under the conditions which probably existed just after the tank rupture. The panel separation shock closed the fuel cell l and 3 oxygen reactant shut-off valves and several propellant and helium isolation valves in the reaction control system. Data were lost for about 1.8 seconds as"
  },
  {
    "chunk_id": 249,
    "question": "What type of valve will be redesigned?",
    "answer": "supply valve",
    "difficulty": "easy",
    "chunk_content": "supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems.\n\nA more thorough discussion of this anomaly is presented in reference l.\n\nThis anomaly is closed.\n\n14.1.2 Postlanding Vent Valve Malfunction\n\nDuring postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen."
  },
  {
    "chunk_id": 249,
    "question": "What type of valve was open during postlanding activities?",
    "answer": "exhaust valve",
    "difficulty": "easy",
    "chunk_content": "supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems.\n\nA more thorough discussion of this anomaly is presented in reference l.\n\nThis anomaly is closed.\n\n14.1.2 Postlanding Vent Valve Malfunction\n\nDuring postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen."
  },
  {
    "chunk_id": 252,
    "question": "What was found to be within specifications in the valve-lock mechanism rigging?",
    "answer": "Tolerances",
    "difficulty": "easy",
    "chunk_content": "The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle.\n\nThe ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found."
  },
  {
    "chunk_id": 252,
    "question": "What happened when the malfunction was duplicated in the spacecraft?",
    "answer": "Only partial travel of the handle",
    "difficulty": "medium",
    "chunk_content": "The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle.\n\nThe ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found."
  },
  {
    "chunk_id": 253,
    "question": "What was the amplitude of the fluctuations observed in the computer readout of the optics shaft angle?",
    "answer": "0.3 degree",
    "difficulty": "easy",
    "chunk_content": "To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan.\n\nThis anomaly is closed.\n\n14.l.3 Shaft Fluctuations in the Zero Optics Mode\n\nBeginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission."
  },
  {
    "chunk_id": 253,
    "question": "What was the system mode when the fluctuations were observed?",
    "answer": "zero optics mode",
    "difficulty": "easy",
    "chunk_content": "To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan.\n\nThis anomaly is closed.\n\n14.l.3 Shaft Fluctuations in the Zero Optics Mode\n\nBeginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission."
  },
  {
    "chunk_id": 257,
    "question": "What type of resolver is used only for the shaft axis?",
    "answer": "Half-speed resolver",
    "difficulty": "easy",
    "chunk_content": "The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation.\n\nSimulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection."
  },
  {
    "chunk_id": 257,
    "question": "What is the equivalent of zero output in shaft rotation?",
    "answer": "Zero degrees",
    "difficulty": "easy",
    "chunk_content": "The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation.\n\nSimulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection."
  },
  {
    "chunk_id": 260,
    "question": "What was the angle in pitch that the Command Module Pilot had manually adjusted the antenna settings to?",
    "answer": "23 degrees",
    "difficulty": "easy",
    "chunk_content": "14.l.4 High-Gain Antenna Acquisition Problem\n\nPrior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station."
  },
  {
    "chunk_id": 260,
    "question": "How far away from the line of sight to the ground station did the antenna boresight axis point due to the difference between the two sets of angles?",
    "answer": "35 degrees",
    "difficulty": "medium",
    "chunk_content": "14.l.4 High-Gain Antenna Acquisition Problem\n\nPrior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station."
  },
  {
    "chunk_id": 261,
    "question": "What was the decrease in uplink signal strength when the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna?",
    "answer": "6 dB",
    "difficulty": "easy",
    "chunk_content": "When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10."
  },
  {
    "chunk_id": 261,
    "question": "In what mode was the high-gain antenna when the reacquisition mode was selected at 55:00:10?",
    "answer": "medium-beam, manual mode",
    "difficulty": "medium",
    "chunk_content": "When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10."
  },
  {
    "chunk_id": 263,
    "question": "What happened to the antenna when it acquired the earth in wide beam?",
    "answer": "The antenna acquired the earth in wide beam.",
    "difficulty": "easy",
    "chunk_content": "limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual"
  },
  {
    "chunk_id": 263,
    "question": "What triggered the switch to manual system?",
    "answer": "the scan limit function line",
    "difficulty": "medium",
    "chunk_content": "limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual"
  },
  {
    "chunk_id": 269,
    "question": "What was the temperature at which the system was cold soaked for 7 hours?",
    "answer": "30\u00b0 F",
    "difficulty": "easy",
    "chunk_content": "An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive.\n\nA complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30\u00b0 F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests.\n\nFollowing verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated."
  },
  {
    "chunk_id": 269,
    "question": "What was the result of the continuous functional tests while the system was warming up?",
    "answer": "The system operated normally",
    "difficulty": "medium",
    "chunk_content": "An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive.\n\nA complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30\u00b0 F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests.\n\nFollowing verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated."
  },
  {
    "chunk_id": 271,
    "question": "What was the temperature of the gas during operation?",
    "answer": "$2000^{\\circ}$ F",
    "difficulty": "easy",
    "chunk_content": "During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).\u3001 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\\circ}$ F\u3002The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units."
  },
  {
    "chunk_id": 271,
    "question": "What was the pressure of the gas during operation?",
    "answer": "14 000 psi",
    "difficulty": "easy",
    "chunk_content": "During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).\u3001 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\\circ}$ F\u3002The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units."
  },
  {
    "chunk_id": 273,
    "question": "What type of material will be applied to the interior of the breech plenum area on future spacecraft?",
    "answer": "Polyimide",
    "difficulty": "easy",
    "chunk_content": "sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition.\n\nThis anomaly is closed.\n\n\n\nFigure. l4-l0.- Tunnel gusset protection.\n\n14.l.7 Reaction Control Isolation Valve Failure\n\nDuring postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation."
  },
  {
    "chunk_id": 273,
    "question": "Where did the miswiring of the fuel valve closing coil occur during initial installation?",
    "answer": "On a terminal board",
    "difficulty": "medium",
    "chunk_content": "sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition.\n\nThis anomaly is closed.\n\n\n\nFigure. l4-l0.- Tunnel gusset protection.\n\n14.l.7 Reaction Control Isolation Valve Failure\n\nDuring postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation."
  },
  {
    "chunk_id": 275,
    "question": "What type of checks are used to verify the circuit in closeout installations?",
    "answer": "Functional checks",
    "difficulty": "easy",
    "chunk_content": "Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired.\n\nThis anomaly is closed.\n\n14.l.8 Potable Water Quantity Fluctuations"
  },
  {
    "chunk_id": 275,
    "question": "What type of testing will be performed on future spacecraft to prove that the isolation valves are properly wired?",
    "answer": "Resistance checks",
    "difficulty": "medium",
    "chunk_content": "Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired.\n\nThis anomaly is closed.\n\n14.l.8 Potable Water Quantity Fluctuations"
  },
  {
    "chunk_id": 276,
    "question": "What was the duration of the first water quantity fluctuation?",
    "answer": "5 minutes",
    "difficulty": "easy",
    "chunk_content": "The potable water quantity measurement fluctuated briefly on two occasions during the mission. At about 23 hours, the reading decreased from 98 to 79 percent for about 5 minutes and then returned to a normal reading of approximately l02 percent. Another fluctuation was noted at about 37 hours, at which time the reading decreased from its upper limit to 83.5 percent. The reading then returned to the upper limit in a period of 7 seconds.\n\nPreflight fluctuations of from 2 to 6 percent near the full level were observed once during the countdown demonstration test, and a possible earlier fluctuation of about 4 percent at the half-full level was noted during the flight readiness test."
  },
  {
    "chunk_id": 276,
    "question": "At what time did the second water quantity fluctuation return to its normal reading?",
    "answer": "in a period of 7 seconds",
    "difficulty": "medium",
    "chunk_content": "The potable water quantity measurement fluctuated briefly on two occasions during the mission. At about 23 hours, the reading decreased from 98 to 79 percent for about 5 minutes and then returned to a normal reading of approximately l02 percent. Another fluctuation was noted at about 37 hours, at which time the reading decreased from its upper limit to 83.5 percent. The reading then returned to the upper limit in a period of 7 seconds.\n\nPreflight fluctuations of from 2 to 6 percent near the full level were observed once during the countdown demonstration test, and a possible earlier fluctuation of about 4 percent at the half-full level was noted during the flight readiness test."
  },
  {
    "chunk_id": 279,
    "question": "At what time did the suit pressure transducer reading suddenly drop from 6.7 to 5.7 psia?",
    "answer": "00:02:45",
    "difficulty": "medium",
    "chunk_content": "During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2).\n\n\n\n(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure."
  },
  {
    "chunk_id": 279,
    "question": "What was the nominal regulated pressure of the cabin?",
    "answer": "5.0 psia",
    "difficulty": "easy",
    "chunk_content": "During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2).\n\n\n\n(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure."
  },
  {
    "chunk_id": 283,
    "question": "What was noted to have escaped from the left-hand electrical circuit interrupter?",
    "answer": "Propellant gas",
    "difficulty": "easy",
    "chunk_content": "During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3."
  },
  {
    "chunk_id": 283,
    "question": "What was found on the adjacent equipment due to the escaped propellant gas?",
    "answer": "Soot",
    "difficulty": "medium",
    "chunk_content": "During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3."
  },
  {
    "chunk_id": 285,
    "question": "What is the percentage of the block that has been crushed when the O-ring enters the chamfer in the breech assembly?",
    "answer": "94",
    "difficulty": "easy",
    "chunk_content": "The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include:\n\na. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block."
  },
  {
    "chunk_id": 285,
    "question": "What is the distance between the piston O-ring and the chamfer in the breech assembly when the fork contacts the attenuator block?",
    "answer": "0.075 inch",
    "difficulty": "medium",
    "chunk_content": "The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include:\n\na. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block."
  },
  {
    "chunk_id": 288,
    "question": "What was the pressure rise rate during the countdown demonstration test?",
    "answer": "7.9 psi/hour",
    "difficulty": "easy",
    "chunk_content": "helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively.\n\nThe heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank."
  },
  {
    "chunk_id": 288,
    "question": "At what pressure range did the helium tank's rise-rate characteristics increase in the manner exhibited during the countdown demonstration test?",
    "answer": "Between 640 and 900 psia",
    "difficulty": "medium",
    "chunk_content": "helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively.\n\nThe heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank."
  },
  {
    "chunk_id": 289,
    "question": "What was the normal prelaunch-standby rise rate of the helium tank?",
    "answer": "7.8 psi/hour",
    "difficulty": "easy",
    "chunk_content": "The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.\u2019 After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard.\n\nThe helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\\mathsf{10}^{-\\gamma}$ torr during the manufacturing process."
  },
  {
    "chunk_id": 289,
    "question": "What material was used for the insulation between the two shells of the helium tank?",
    "answer": "aluminized Mylar",
    "difficulty": "medium",
    "chunk_content": "The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.\u2019 After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard.\n\nThe helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\\mathsf{10}^{-\\gamma}$ torr during the manufacturing process."
  },
  {
    "chunk_id": 290,
    "question": "What is the likely cause of the anomaly in the tank-insulation?",
    "answer": "Tank-insulation degradation",
    "difficulty": "medium",
    "chunk_content": "The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during"
  },
  {
    "chunk_id": 290,
    "question": "What happens to the contaminant when it freezes?",
    "answer": "The pressure of the contaminant is too low to significantly affect the thermal conductivity.",
    "difficulty": "easy",
    "chunk_content": "The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during"
  },
  {
    "chunk_id": 292,
    "question": "What is the desired steady-state pressure rise rate for a perfect tank over the entire range of temperatures?",
    "answer": "8 psi/hour",
    "difficulty": "easy",
    "chunk_content": "Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure.\n\nA screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\\circ}$ to $\\bar{\\mathsf{1}}\\bar{2}\\bar{3}^{\\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity.\n\nThis anomaly is closed.\n\n14.2.2 Abnormal Descent Stage Noise"
  },
  {
    "chunk_id": 292,
    "question": "How long after pumpdown is the pressure in the jacket measured to verify vacuum integrity?",
    "answer": "2 or 3 weeks",
    "difficulty": "medium",
    "chunk_content": "Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure.\n\nA screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\\circ}$ to $\\bar{\\mathsf{1}}\\bar{2}\\bar{3}^{\\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity.\n\nThis anomaly is closed.\n\n14.2.2 Abnormal Descent Stage Noise"
  },
  {
    "chunk_id": 293,
    "question": "What was the duration of the current transients in the descent batteries?",
    "answer": "2 seconds",
    "difficulty": "easy",
    "chunk_content": "At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down.\n\n\n\nFigure 14-l5.- Descent stage battery location.\n\n\n\nThe thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition."
  },
  {
    "chunk_id": 293,
    "question": "Where was the momentary short circuit located in the dc electrical system?",
    "answer": "Lunar-Module-Pilot side",
    "difficulty": "medium",
    "chunk_content": "At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down.\n\n\n\nFigure 14-l5.- Descent stage battery location.\n\n\n\nThe thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition."
  },
  {
    "chunk_id": 294,
    "question": "What percentage of the total current load was battery 2 providing immediately after the current surges?",
    "answer": "80 percent",
    "difficulty": "easy",
    "chunk_content": "The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section."
  },
  {
    "chunk_id": 294,
    "question": "At what time did battery 2 give an indication of a battery malfunction?",
    "answer": "99:5l:09",
    "difficulty": "medium",
    "chunk_content": "The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section."
  },
  {
    "chunk_id": 301,
    "question": "What was the maximum differential pressure that caused the leak in the oxygen tank 2?",
    "answer": "193 psi",
    "difficulty": "easy",
    "chunk_content": "During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction.\n\nThe internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring\n\n\n\nFigure 14-l9.- Oxygen-supply system."
  },
  {
    "chunk_id": 301,
    "question": "How much leakage was allowed for the valve in either direction?",
    "answer": "360 scc/hr",
    "difficulty": "easy",
    "chunk_content": "During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction.\n\nThe internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring\n\n\n\nFigure 14-l9.- Oxygen-supply system."
  },
  {
    "chunk_id": 302,
    "question": "What is the recommended compression range for the O-ring by the manufacturer?",
    "answer": "between 0.01l5 and 0.0225 inch",
    "difficulty": "medium",
    "chunk_content": "(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination.\n\nThe production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly.\n\n\n\nFigure l4-20.- Ascent stage tank shutoff valve:."
  },
  {
    "chunk_id": 302,
    "question": "What is the cause of the leakage in the other two valves?",
    "answer": "contamination",
    "difficulty": "easy",
    "chunk_content": "(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination.\n\nThe production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly.\n\n\n\nFigure l4-20.- Ascent stage tank shutoff valve:."
  },
  {
    "chunk_id": 303,
    "question": "What type of material is stitched to the inner surface of the window shade?",
    "answer": "A Beta Cloth",
    "difficulty": "easy",
    "chunk_content": "Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves.\n\nThis anomaly is closed.\n\n14.2.5 Cracked Window Shade\n\nThe left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade.\n\n\n\nFigure 14-21.- Cracked left-hand window shade.\n\nCracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight."
  },
  {
    "chunk_id": 303,
    "question": "What percentage of stitch holes had cracks extending from them on the window shade?",
    "answer": "80 percent",
    "difficulty": "medium",
    "chunk_content": "Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves.\n\nThis anomaly is closed.\n\n14.2.5 Cracked Window Shade\n\nThe left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade.\n\n\n\nFigure 14-21.- Cracked left-hand window shade.\n\nCracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight."
  },
  {
    "chunk_id": 304,
    "question": "What percentage of elongation does the modified Aclar material provide?",
    "answer": "25 percent",
    "difficulty": "easy",
    "chunk_content": "The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight.\n\nThis anomaly is closed.\n\n14.3 GOVERNMENT FURNISHED EQUIPMENT\n\n14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera\n\nFor launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies."
  },
  {
    "chunk_id": 304,
    "question": "What type of material will the shades for future vehicles be fabricated from?",
    "answer": "Aclar",
    "difficulty": "easy",
    "chunk_content": "The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight.\n\nThis anomaly is closed.\n\n14.3 GOVERNMENT FURNISHED EQUIPMENT\n\n14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera\n\nFor launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies."
  },
  {
    "chunk_id": 305,
    "question": "What was used to secure the set screw of the interval timer set knob prior to the change?",
    "answer": "A special gripping compound",
    "difficulty": "easy",
    "chunk_content": "To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly.\n\nThis anomaly is closed.\n\n14.3.2 Failure of the Interval Timer Set Knob\n\nThe onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin."
  },
  {
    "chunk_id": 305,
    "question": "What will be used to secure the knobs on timers for future flights?",
    "answer": "A roll pin",
    "difficulty": "easy",
    "chunk_content": "To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly.\n\nThis anomaly is closed.\n\n14.3.2 Failure of the Interval Timer Set Knob\n\nThe onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin."
  },
  {
    "chunk_id": 307,
    "question": "What type of drops will be packaged the same as eye drops in medical kits for future flights?",
    "answer": "Nose drops",
    "difficulty": "easy",
    "chunk_content": "Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops.\n\nThis anomaly is closed.\n\n15.0 CONCLUS IONS\n\nThe Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report."
  },
  {
    "chunk_id": 307,
    "question": "What was the Apollo mission that required an emergency abort?",
    "answer": "Apollo l3",
    "difficulty": "medium",
    "chunk_content": "Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops.\n\nThis anomaly is closed.\n\n15.0 CONCLUS IONS\n\nThe Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report."
  },
  {
    "chunk_id": 309,
    "question": "What was the purpose of the lunar flyby mission?",
    "answer": "a lunar flyby mission, including three planned experiments",
    "difficulty": "easy",
    "chunk_content": "d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency.\n\ne. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived.\n\nThe configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented."
  },
  {
    "chunk_id": 309,
    "question": "What was the outcome of the mission regarding the lunar module's backup capability?",
    "answer": "information which would have otherwise been unavailable",
    "difficulty": "medium",
    "chunk_content": "d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency.\n\ne. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived.\n\nThe configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented."
  },
  {
    "chunk_id": 315,
    "question": "What was the original location of the sensing point for the water separator drain tank?",
    "answer": "The carbon dioxide sensor",
    "difficulty": "easy",
    "chunk_content": "propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush"
  },
  {
    "chunk_id": 315,
    "question": "What was the purpose of adding a removable flow limiter to the inlet for the primary lithium hydroxide cartridge?",
    "answer": "To reduce the water separator speed and to minimize the possibility of condensed water in the suit.",
    "difficulty": "medium",
    "chunk_content": "propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush"
  },
  {
    "chunk_id": 316,
    "question": "What material were the bristles on the vacuum brush changed from?",
    "answer": "Teflon",
    "difficulty": "easy",
    "chunk_content": "was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag."
  },
  {
    "chunk_id": 316,
    "question": "What was added to the lunar sample tote bag?",
    "answer": "a cover",
    "difficulty": "easy",
    "chunk_content": "was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag."
  },
  {
    "chunk_id": 318,
    "question": "What instruments were deleted from the Apollo 13 package compared to Apollo 12?",
    "answer": "Solar wind spectrometer, magnetometer, and suprathermal ion detector",
    "difficulty": "easy",
    "chunk_content": "The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13.\n\nThe Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay.\n\nNASA-S-70-5864\n\n\n\nFigure A-l.- Experiment subpackage number l.\n\nA.3.1 Heat Flow Experiment\n\nThe heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements."
  },
  {
    "chunk_id": 318,
    "question": "What was the purpose of the heat flow experiment on Apollo 13?",
    "answer": "To measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials",
    "difficulty": "medium",
    "chunk_content": "The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13.\n\nThe Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay.\n\nNASA-S-70-5864\n\n\n\nFigure A-l.- Experiment subpackage number l.\n\nA.3.1 Heat Flow Experiment\n\nThe heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements."
  },
  {
    "chunk_id": 321,
    "question": "What type of insulation was used in the S-Il stage?",
    "answer": "Spray foam",
    "difficulty": "easy",
    "chunk_content": "The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover.\n\nA.4 LAUNCH VEHICLE\n\nSpray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation.\n\nA.5 MASS PROPERTIES"
  },
  {
    "chunk_id": 321,
    "question": "Where were telemetry measurements relocated to provide a more complete analysis of platform vibrations?",
    "answer": "In the inertial platform",
    "difficulty": "medium",
    "chunk_content": "The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover.\n\nA.4 LAUNCH VEHICLE\n\nSpray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation.\n\nA.5 MASS PROPERTIES"
  },
  {
    "chunk_id": 323,
    "question": "What is the moment of inertia of the Command&service modules?",
    "answer": "847.4",
    "difficulty": "easy",
    "chunk_content": "Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth"
  },
  {
    "chunk_id": 323,
    "question": "What is the product of inertia of the Lunar module before the cryotenic oxygen tank ignition?",
    "answer": "33995",
    "difficulty": "medium",
    "chunk_content": "Event Weight, 1b Center of gravity, in. Moment or inertia, slug-ft2 Product of inertia, slug-ft2 X Z IxY Lift-off 110 252.4 847.4 2.4 3.7 67646 1 175 539 1 178 016 2906 8047 3711 Earth orbit insertion 101 261.2 807.4 2.6 4.1 66770 718 686 721 213 5157 11945 3688 Command&servicemodules Lwnar module 63 720.3 33499.1 934.5 1237.0 4.0 -0.1 6.5 0.0 33995 22457 76486 24654 79123 25255 ~1746 -126 95 3221 235 Totaldocked 97 219.4 1038.7 2.6 4.3 56 736 534890 538009 -8142 -9376 3585 First midcourse correction Ignition Cutoff 97 081.5 96 851.1 1038.9 1039.0 2.6 2.6 4.2 4.2 56 629 56 508 534493 534 139 537 635 537 380 -8192 -8189 -9305 -9282 3620 3587 Cryotenic oxygen tank incitent Before 96 646.9 1039.2 2.6 4.2 56 321 533499 536 766 -8239 -9244 3636 After Second midcourse correction 96 038.7 1040.7 3.0 3.9 57248 533 927 537 251 -8269 669- -3709 Ignition Cutoff 95 959.9 95 647.1 378.8 379.4 4.9 5.0 0.7 0.7 57205 57006 516443 513919 521 180 518700 11617 11553 2659 2651 3286 3285 Transearth"
  },
  {
    "chunk_id": 324,
    "question": "What was the ignition cutoff for the third midcourse correction?",
    "answer": "398.7",
    "difficulty": "easy",
    "chunk_content": "injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8"
  },
  {
    "chunk_id": 324,
    "question": "What was the altitude of the command module after lunar module separation?",
    "answer": "2362",
    "difficulty": "medium",
    "chunk_content": "injection Ignition Cutoff 95 424.0 87456.0 379.7 398.4 5.0 5.5 0.7 0.8 56 866 51778 512 837 431285 517 560 437119 11370 9443 2495 2222 3255 3249 Thirdmidcourse correction Ignition Cutoff 87 325.3 87 263.3 398.7 398.9 5.5 5.5. 0.8 0.8 51 681 51642 430 123 429353 435930 435 169 9244 9227 2048 3215 Fowrth midcourse correction Ignition 87 132.1 399.1 5.5 0.8 51 553 428 322 434105 9069 2045 1911 3215 3191 Cutoff Command&servicemodule 87 101.5 399.2 5.6 0.8 51538 428219 433990 9065 1910 3192 b separation Before 87 057.3 399.3 5.6 0.8 51 517 428065 433 819 9058 1909 3194 After (command module/ lunar module) 37 109.7 251.5 2.2 -0.3 24048 92418 93.809 2362 686= 9 Commandmodule/lunar module separation b 37 014.6 Before 12 367.6 252.9 1039.9 1.9 -0.6 23926 93993 95514 2188 -963 -35 After (command module) 0.3 6.1 581.5 5 258 4636 31 409 20 Entry 12 361.4 1039.9 0.3 6.0 5 812 5254 4635 31 -407 21 Drogue deployment 11 869.4 1038.7 0.3 6.0 5727 5002 4405 33 -382 24 Main parachute deployment 11 579.8"
  },
  {
    "chunk_id": 325,
    "question": "What is the first number in the list?",
    "answer": "1038.6",
    "difficulty": "easy",
    "chunk_content": "1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42"
  },
  {
    "chunk_id": 325,
    "question": "What is the last word in the list?",
    "answer": "Landing",
    "difficulty": "easy",
    "chunk_content": "1038.6 0.5 5.3 5590 4 812 4346 27 -319 41 Landing 11 132.9 1036.6 0.5 5.2 5526 4531 4046 25 -328 42"
  },
  {
    "chunk_id": 326,
    "question": "Where was the history of command and service module (cSM 109) operations at the manufacturer's facility shown?",
    "answer": "Downey, California",
    "difficulty": "easy",
    "chunk_content": "ALunar module was docked to the command module from initial docking wntil just prior to entry. \"Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases.\n\nThe history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2.\n\nThe history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4.\n\n\n\nFigure B-l.- Checkout flow for command and service modules at contractor's facility.\n\nNASA-S-70-5867\n\n\n\nFigure B-2.- Command and service module checkout history at Kennedy Space Center.\n\nNASA-S-70-5868\n\n\n\nFigure B-3.- Checkout flow for lunar module at contractor's facility.\n\n\n\nFigure $\\mathbb{R}{-}\\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center."
  },
  {
    "chunk_id": 326,
    "question": "Where was the history of the lunar module (LM-7) operations at Kennedy Space Center shown?",
    "answer": "Figure B-4",
    "difficulty": "medium",
    "chunk_content": "ALunar module was docked to the command module from initial docking wntil just prior to entry. \"Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases.\n\nThe history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2.\n\nThe history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4.\n\n\n\nFigure B-l.- Checkout flow for command and service modules at contractor's facility.\n\nNASA-S-70-5867\n\n\n\nFigure B-2.- Command and service module checkout history at Kennedy Space Center.\n\nNASA-S-70-5868\n\n\n\nFigure B-3.- Checkout flow for lunar module at contractor's facility.\n\n\n\nFigure $\\mathbb{R}{-}\\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center."
  },
  {
    "chunk_id": 332,
    "question": "What is the time range of the first recorded data point?",
    "answer": "57:57",
    "difficulty": "easy",
    "chunk_content": "Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X"
  },
  {
    "chunk_id": 332,
    "question": "What is the type of data recorded in the last entry of the table?",
    "answer": "MSFN",
    "difficulty": "medium",
    "chunk_content": "Time, hr:min Range station Bandpass plots or tabs Bilevels Computer spxOM O'graph records Brush records Special plots or tabs Special programs From To 57:57 57:57 58:12 60:36 61:10 64:52 65:07 68:26 72:32 77:03 78:47 80:29 93:30 94:56 96:29 97:11 97:12 99:24 99:50 100:33 101:00 104 :19 104:57 105 :15 108:36 108:52 109 :12 112:35 117 :33 102:28 133:46 58:05 60:36 59:12 64:50 62:10 68:26 66:07 72:24 77:03 80:29 79:47 96:29 93:40 95 :05 100:33 97:18 97:18 99:57 100:24 104:57 101:07 105:19 108:36 105:53 112: 35 109:08 109:25 GDS MSFN GDS MSFN GDS MSFN HSK MSFN MSFN MSFN SGD MSFN MAD MAD N&SW MAD ACN GDS GDS MSFN GDS GDS MSFN GDS MSFN GDS GDS 120:28 MSFN MAD 136:52 MSFN 134:39 GDS HSK 136:50 HSK 138:34 MSFN 138:14 HSK 142:38 MSFN 141:35 CRO X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X X"
  },
  {
    "chunk_id": 334,
    "question": "What was the publication date of the report titled \"Reaction Control System Performance\"?",
    "answer": "August 1969",
    "difficulty": "easy",
    "chunk_content": "Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command"
  },
  {
    "chunk_id": 334,
    "question": "How many reports were published for the Apollo 8 mission?",
    "answer": "7",
    "difficulty": "medium",
    "chunk_content": "Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command"
  },
  {
    "chunk_id": 335,
    "question": "What was the year the Ascent Propulsion System received its Final Flight Evaluation?",
    "answer": "1969",
    "difficulty": "easy",
    "chunk_content": "and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969"
  },
  {
    "chunk_id": 335,
    "question": "What was the status of the item listed as number 9?",
    "answer": "Cancelled",
    "difficulty": "easy",
    "chunk_content": "and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969"
  },
  {
    "chunk_id": 336,
    "question": "What was the publication date of the report titled \"Performance Analysis\"?",
    "answer": "December 1969",
    "difficulty": "easy",
    "chunk_content": "Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo \uff0911 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review"
  },
  {
    "chunk_id": 336,
    "question": "What was the status of the report titled \"Analysis of Apollo lo Photography and Visual Observations\"?",
    "answer": "In publication",
    "difficulty": "medium",
    "chunk_content": "Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo \uff0911 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review"
  },
  {
    "chunk_id": 337,
    "question": "What was the title of the report about the Apollo 12 trajectory?",
    "answer": "Apollo 12 Trajectory Reconstruction and Analysis",
    "difficulty": "easy",
    "chunk_content": "Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation\n\nREFERENCES\n\nManned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970.\n\nMarshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970."
  },
  {
    "chunk_id": 337,
    "question": "What was the status of the report \"Apollo 12 Trajectory Reconstruction and Analysis\" when it was published?",
    "answer": "1",
    "difficulty": "easy",
    "chunk_content": "Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation\n\nREFERENCES\n\nManned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970.\n\nMarshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970."
  },
  {
    "chunk_id": 338,
    "question": "What year was the Study Conference on the Global Atmospheric Research Program held?",
    "answer": "1967",
    "difficulty": "easy",
    "chunk_content": "Marshall Space Flight Center, Kennedy Space Center, Manned Spacecraft Center: Analysis of Apollo l2 Lightning Incident, MSC-01540. February 1970.\n\nICSU/IUGG Committee on Atmospheric Sciences: Report of the Study Conference on the Global Atmospheric Research Program, 1967.\n\nBulletin of the American Meteorological Society, Vol. 50, No. 7: Cloud Height Contouring from Apollo 6 Photography, by V. S. Whitehead, I. D. Browne, and J. G. Garcia. 1969.\n\nDefense Supply Agency, Washington, D. C.: Military Standardization Handbook_ Optical Design, MIL HDBK-14l. 1962.\n\nNASA Headquarters: Apollo Flight Mission Assignments. OMSF M-D MA500-11 (SE 010-000-1). 0ctober 1969.\n\nManned Spacecraft Center: Mission Requirement, H-2 Type Mission (Lunar Landing). SPD9-R-053. November 10, 1969.\n\nAPOLLO SPACECRAFT FLIGHT HISTORY\n\n(Continued from inside front cover)"
  },
  {
    "chunk_id": 338,
    "question": "Who authored the article \"Cloud Height Contouring from Apollo 6 Photography\"?",
    "answer": "V. S. Whitehead, I. D. Browne, and J. G. Garcia",
    "difficulty": "medium",
    "chunk_content": "Marshall Space Flight Center, Kennedy Space Center, Manned Spacecraft Center: Analysis of Apollo l2 Lightning Incident, MSC-01540. February 1970.\n\nICSU/IUGG Committee on Atmospheric Sciences: Report of the Study Conference on the Global Atmospheric Research Program, 1967.\n\nBulletin of the American Meteorological Society, Vol. 50, No. 7: Cloud Height Contouring from Apollo 6 Photography, by V. S. Whitehead, I. D. Browne, and J. G. Garcia. 1969.\n\nDefense Supply Agency, Washington, D. C.: Military Standardization Handbook_ Optical Design, MIL HDBK-14l. 1962.\n\nNASA Headquarters: Apollo Flight Mission Assignments. OMSF M-D MA500-11 (SE 010-000-1). 0ctober 1969.\n\nManned Spacecraft Center: Mission Requirement, H-2 Type Mission (Lunar Landing). SPD9-R-053. November 10, 1969.\n\nAPOLLO SPACECRAFT FLIGHT HISTORY\n\n(Continued from inside front cover)"
  },
  {
    "chunk_id": 339,
    "question": "What was the launch date of the Apollo 4 mission?",
    "answer": "Nov. 9, 1967",
    "difficulty": "easy",
    "chunk_content": "Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11\uff0c1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18"
  },
  {
    "chunk_id": 339,
    "question": "What was the purpose of the Apollo 6 mission?",
    "answer": "Verification of closed-loop emergency detection system",
    "difficulty": "medium",
    "chunk_content": "Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11\uff0c1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18"
  },
  {
    "chunk_id": 340,
    "question": "What is the location of the NASA-Manned Spacecraft Center?",
    "answer": "Houston, Texas 77058",
    "difficulty": "easy",
    "chunk_content": "POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION\n\nNASA-Manned Spacecraft Center Houston, Texas 77058\n\nATTN: PT2(office Symbol)\n\n(Continued from inside front cover)"
  },
  {
    "chunk_id": 340,
    "question": "What is the recipient of the letter?",
    "answer": "PT2",
    "difficulty": "easy",
    "chunk_content": "POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION\n\nNASA-Manned Spacecraft Center Houston, Texas 77058\n\nATTN: PT2(office Symbol)\n\n(Continued from inside front cover)"
  },
  {
    "chunk_id": 341,
    "question": "What was the date of the first lunar landing?",
    "answer": "July 16, 1969",
    "difficulty": "easy",
    "chunk_content": "Mi ssion Spacecraft Description Launch date Launch site Apollo4 SC-017 LTA-10R Supercircular entry at lunar Nov.9,1967 Kennedy Space Center, Fla. Apollo 5 LM-1 return velocity First lunar module flight Jan.22,1968 Cape Kennedy, Apollo 6 SC-020 LTA-2R Verification of closed-loop April 4, 1968 Fla. Kennedy Space Center, Fla. Apollo7 CSM 101 emergency detection system First manned flight; Oct.11\uff0c1968 Apol1o 8 CSM 103 earth-orbital First manned lunar Dec.2l,1968 Cape Kennedy, Fla. Kennedy Space Apol1o9 CSM 104 orbital flight; first manned Saturn V launch First manned lunar Apollo 10 LM-3 module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center, Fla. CSM 106 LM-4 First lunar orbit rendezvous; low pass over lumar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 Second lunar landing Center, Fla. Apollo 13 LM-6 Nov. 14, 1969 Kennedy Space Center, Fla."
  },
  {
    "chunk_id": 341,
    "question": "What was the name of the launch site for Apollo 6?",
    "answer": "Kennedy Space Center, Fla.",
    "difficulty": "medium",
    "chunk_content": "Mi ssion Spacecraft Description Launch date Launch site Apollo4 SC-017 LTA-10R Supercircular entry at lunar Nov.9,1967 Kennedy Space Center, Fla. Apollo 5 LM-1 return velocity First lunar module flight Jan.22,1968 Cape Kennedy, Apollo 6 SC-020 LTA-2R Verification of closed-loop April 4, 1968 Fla. Kennedy Space Center, Fla. Apollo7 CSM 101 emergency detection system First manned flight; Oct.11\uff0c1968 Apol1o 8 CSM 103 earth-orbital First manned lunar Dec.2l,1968 Cape Kennedy, Fla. Kennedy Space Apol1o9 CSM 104 orbital flight; first manned Saturn V launch First manned lunar Apollo 10 LM-3 module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center, Fla. CSM 106 LM-4 First lunar orbit rendezvous; low pass over lumar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 Second lunar landing Center, Fla. Apollo 13 LM-6 Nov. 14, 1969 Kennedy Space Center, Fla."
  }
]