Question: What was the nominal ground pressure rise rate during the countdown demonstration test?

Answer: 7.8 psi/hr

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.

The 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.


Question: What was the average rise rate from lift-off to the first descent propulsion maneuver?

Answer: 7.0 psi/hr

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.

The 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.


Question: What was the predicted rupture range for the vehicle's helium tank?

Answer: 190o ± 20 psia

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 ± 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


Question: What shape are the holes ported on two sides of the vent tube for the supercritical helium tank?

Answer: Oval-shaped

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 ± 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


Question: How many hours did the environmental control system provide a habitable environment for the crew?

Answer: 83

Environmental. 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.

An 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.


Question: What was the average water usage rate at the higher end of the range?

Answer: 6.3 lb/hr

Environmental. 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.

An 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.


Question: What was partially restricted with tape in the suit-circuit arrangement?

Answer: The mass flow

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.


Question: How long was the lunar module suit circuit operating before an additional unit was stacked on each original cartridge?

Answer: Approximately 20 hours

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.


Question: What was the total amount of propellant loaded into the tanks for the service propulsion system?

Answer: 25084

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.

The 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.

Fuel, 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


Question: What was the amount of fuel consumed by the service propulsion system?

Answer: 92.3

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.

The 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.

Fuel, 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


Question: What was the total oxygen supply in the surge tank after the incident?

Answer: 3.77 pounds

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

Following 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.

At 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.


Question: How much water was transferred from the command module to the lunar module during the abort phase?

Answer: 14 pounds

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

Following 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.

At 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.


Question: What was the amount of energy remaining in the batteries at landing?

Answer: 29 ampere-hours

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.

Following 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.

The 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.


Question: What was the original amount of energy remaining in the three entry batteries?

Answer: 99 ampere-hours

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.

Following 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.

The 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.


Question: What was the total amount of fuel loaded in System A?

Answer: 316.5

The 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.

Fuel, lb Oxidi zer， 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

Actual 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.

Loaded; 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


Question: How much oxygen was consumed by the Ascent stage Tank 2?

Answer: 2.4

The 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.

Fuel, lb Oxidi zer， 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

Actual 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.

Loaded; 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


Question: What was the total loaded water available for inflight use at launch?

Answer: 338 pounds

arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4).

During 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·is shown in figure 7.2-l.


Question: How many hours of cooling would 50 pounds of water provide at the reduced power condition?

Answer: 18 hours

arhe shutoff valve in ascent stage tank 2 had reverse leakage (discussed in section 14.2.4).

During 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·is shown in figure 7.2-l.


Question: What was the initial current consumption of the vehicle before the second descent propulsion system firing?

Answer: 30 amperes

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

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.

Figure 7.2-2.- Lunar module total battery capacity during flight.


Question: How many ampere hours remained in the lunar module batteries at the time of undocking?

Answer: 410

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

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.

Figure 7.2-2.- Lunar module total battery capacity during flight.


Question: What was the name of the Lunar Module Pilot mentioned in the text?

Answer: Fred W. Haise, Jr.

Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr.


Question: Who was the Command Module Pilot mentioned in the text?

Answer: John L. Swigert, Jr.

Commander James A. Lovell, Jr., Command Module Pilot John L. Swigert, Jr., and Lunar Module Pilot Fred W. Haise, Jr.


Question: What was the location of the crew's training base until December 1969?

Answer: Houston

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


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

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


Question: What was the time at which S-II staging and S-IVB ignition occurred?

Answer: 9 minutes 57 seconds

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


Question: How long after S-II staging did engine 5 shut down?

Answer: 2 minutes

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


Question: What was the time of S-IVB engine cutoff?

Answer: 00:12:30

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.


Question: What was the velocity of the spacecraft at insertion?

Answer: 25 565 ft/sec

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.


Question: What was the range of hours during which the S-IVB vibration was experienced?

Answer: 69 to 122 hours

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 ,

(c) 69 to 122 hours. Figure 8-l.- Continued.

(a) 122 to 143 hours. Figure 8-l.- Concluded.

low-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.


Question: What was the computer-displayed inertial velocity at cutoff?

Answer: 35 560 ft/sec

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 ,

(c) 69 to 122 hours. Figure 8-l.- Continued.

(a) 122 to 143 hours. Figure 8-l.- Concluded.

low-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.


Question: What was the rate of the manual pitch maneuver executed after separation and translation?

Answer: 1.5 deg/sec

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


Question: How much reaction control fuel was used for transposition, docking, and extraction?

Answer: 55 pounds

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


Question: How long did earth weather photography last?

Answer: approximately 6 hours

Following translunar injection, earth weather photography was conducted for approximately 6 hours.

The 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.


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

Following translunar injection, earth weather photography was conducted for approximately 6 hours.

The 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.


Question: What roll rate was used with the positive longitudinal spacecraft axis pointed toward ecliptic north pole?

Answer: 0.3 deg/sec

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

alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed.

At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground.


Question: What condition was confirmed by the ground at about 47 hours?

Answer: Oxygen tank 2 quantity sensor failed full scale high

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

alignments (Program 52) with passive thermal control mode rates of 0.3 deg/sec were satisfactory in the optics resolve mode at medium speed.

At about 47 hours the oxygen tank 2 quantity sensor failed full scale high, a condition which was confirmed by the ground.


Question: What was the duration of the planned descent propulsion system maneuver?

Answer: 6l-l/2 hours

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."

Primary 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.


Question: What system was used to maintain attitude after the error needles were nulled?

Answer: primary guidance

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."

Primary 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.


Question: What was the initial throttle position?

Answer: idle position

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.


Question: What was the final throttle position?

Answer: full throttle

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.


Question: How long did it take to fully charge battery B?

Answer: Approximately 3 hours

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.


Question: What was the source of power for the command module during the transearth coast?

Answer: The lunar module through the umbilical connectors

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.


Question: What was the result of the sextant star check and moon occultation?

Answer: Confirmed

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.


Question: What was the first acceleration peak during the guided entry?

Answer: Approximately 5g

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.


Question: What type of flotation equipment is preferable for ease of donning and egress?

Answer: Standard Navy life vests

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.

This 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.

9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA


Question: How long did the first 2 days of the Apollo 13 mission remain routine from a medical point of view?

Answer: Completely

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.

This 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.

9.1 BIOINSTRUMENTATION AND PHYSIOLOGICAL DATA


Question: What was the heart rate of the Lunar Module Pilot?

Answer: 72

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.

Prior 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.

Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12


Question: How many biomedical signals could be monitored in the Lunar Module at a time?

Answer: one

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.

Prior 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.

Crewman Heart rate, beats/min Respiratory rate, breaths/min Commander 68 18 Command Module Pilot 65 15 Lumar Module Pilot 72 12


Question: What was the range of the Command Module Pilot's heart rate during the entry phase?

Answer: 60 to 70 beats/min

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.

During 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.


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

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.

During 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.


Question: What was the nickel content in the command module hot water port?

Answer: $0.05~\mathrm{mg/1}$

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.


Question: What was the outcome of the bacterial population after inflight chlorination?

Answer: to specification levels

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.


Question: How many ounces of water were consumed by each crewman after the oxygen tank incident?

Answer: 24

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.

The 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.


Question: What was the basis for selecting the flight menus?

Answer: Crew preferences determined by preflight evaluation of representative flight foods

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.

The 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.


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."

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.

Preflight 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.


Question: How many physical examinations were conducted for the primary crew before launch?

Answer: 3

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.

Preflight 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.


Question: How many days before flight did the primary Command Module Pilot get exposed to rubella?

Answer: Eight

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.


Question: What was the result of the laboratory studies on the primary Command Module Pilot?

Answer: No immunity to rubella

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.


Question: What was the weight loss of the Lunar Module Pilot?

Answer: 6.5 pounds

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.

The 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.


Question: What was the cause of the Lunar Module Pilot's dizziness?

Answer: Fatigue, the effects of weightlessness, and the urinary tract infection

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.

The 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.


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.

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:

a. 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


Question: What system was used for life support after the oxygen tanks experienced a problem?

Answer: Lunar module

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:

a. 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


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

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


Question: What type of maneuver was used to expedite the landing to about 142:30:00?

Answer: pericynthion-plus-2-hour maneuver (transearth injection)

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


Question: What system was used by the crew to remove carbon dioxide in the lunar module?

Answer: Environmental control system

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.


Question: What was established using the lunar module reaction control system?

Answer: Passive thermal control mode

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.


Question: What was the task that was verified in a simulator before advising the crew?

Answer: procedures

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.


Question: How long before entry did lunar module undocking occur?

Answer: 1 hour

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.


Question: What type of aircraft is capable of lifting the command module?

Answer: HH-53C Helicopters

Landing 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, ’Azores 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

&arotal ship support $=5$ Total aircraft support $\approx$ 23

10.3.l Command Module Location and Retrieval


Question: How many SH-3 Helicopters are staged from Norfolk NAS, Virginia?

Answer: 1

Landing 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, ’Azores 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

&arotal ship support $=5$ Total aircraft support $\approx$ 23

10.3.l Command Module Location and Retrieval


Question: Where was the flight crew flown to after being flown from Pago Pago, Samoa?

Answer: Hawaii

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.

Upon 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.

The following is a chronological listing of events during the recovery operations.


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

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.

Upon 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.

The following is a chronological listing of events during the recovery operations.


Question: When was the flotation collar inflated?

Answer: 1824

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


Question: What was completed at 1936 on April 18?

Answer: Flight crew departed Iwo Jima

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


Question: What was the condition of the interior surfaces of the command module?

Answer: very damp and cold

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.


Question: What was the state of the postlanding ventilation valve unlock handle?

Answer: jammed between the lock and unlock positions

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.


Question: How many electric field meters were installed in the area to the north and west of the launch site?

Answer: Nine

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.


Question: What was the purpose of the special device operated at site 5?

Answer: To measure rapid changes in the electric field

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.


Question: What was the nature of the initial change in the record at site 6?

Answer: Positive

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

Figure ll.l-l.- Field meter location in the laumch site area.

Figure ll.l-2.-- Field meter locations in the proximity of the launch complex.

gravel 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.


Question: What was the approximate negative field value at site 6 after 40 seconds?

Answer: 3000 volts/meter

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

Figure ll.l-l.- Field meter location in the laumch site area.

Figure ll.l-2.-- Field meter locations in the proximity of the launch complex.

gravel 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.


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

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.

Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch.

No 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.


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

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.

Figure ll.l-3.- Electrical discharge data for the Apollo 13 launch.

No 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.


Question: Did the field-change and sferics detectors at site 5 detect any lightning-like discharge during the launch?

Answer: No

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.


Question: What was the origin of the sporadic signals recorded during the afternoon of launch day?

Answer: Lightning in a cold front

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.


Question: What is the equilibrium potential for a conventional jet aircraft?

Answer: a million volts

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.

ll.l.2 Very-Low and Low-Frequency Radio Noise


Question: What is the estimated electrostatic potential of a Saturn V?

Answer: several million volts

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.

ll.l.2 Very-Low and Low-Frequency Radio Noise


Question: What was the frequency to which the first receiver was tuned?

Answer: 1.5 kHz

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.

During 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.


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

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.

During 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.


Question: At what time were the first two balloons launched?

Answer: April 9, 1970

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

altitude 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.

11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES


Question: What was the frequency of the oscillating current before the marked increase in amplitude?

Answer: 15 cycles per minute

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

altitude 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.

11.2 EARTH PHOTOGRAPHY APPLIED TO GEOSYNCHRONOUS SATELLITES


Question: How many photographs were taken for the test?

Answer: Eleven

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,


Question: What was required to reconstruct the geometry involved in the test?

Answer: A precise record of time of photography

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,


Question: What is the altitude of the object in the 13-60-8594 frame?

Answer: 1982.8

TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY


Question: What is the latitude of the object in the 13-60-8591 frame?

Answer: 28°25'N

TABIE 11.2-I.- EARTH WEATHER PHOTOGRAPHY


Question: How long did seismic signals continue after impact?

Answer: over 4 hours

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.


Question: What was the peak signal intensity compared to the Apollo 12 ascent stage impact?

Answer: 8 times larger

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.


Question: What is the estimated depth range of the S-IVB seismic energy in the moon?

Answer: 20 to 40 kilometers

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.


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

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.


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.

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.

Thirteen 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:

TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS


Question: How many detailed objectives were derived from the four primary objectives?

Answer: Thirteen

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.

Thirteen 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:

TABLE 12-I.- DETAILED OBJECTIVES AND EXPERIMENTS


Question: What was the purpose of the S-059 experiment?

Answer: Lunar field geology

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


Question: Was the pilot involved in describing a function during the mission?

Answer: Yes

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


Question: What was the initial azimuth of the vehicle launch?

Answer: 90 degrees east of north

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


Question: How many seconds later than predicted did the event of orbital insertion occur?

Answer: 44.07

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


Question: What was the percentage of the maximum lateral loads experienced during S-IC boost compared to the design value?

Answer: 25 percent

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 ±225 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


Question: At what time did the S-II crossbeam oscillations reach a peak amplitude of +33.7g?

Answer: 330.6 seconds

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 ±225 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


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

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.

This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission.

14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure

At 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:


Question: What was the condition associated with the master alarm that occurred during the Apollo 13 mission?

Answer: Main-bus-B undervoltage condition

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.

This section contains a discussion of the significant problems or discrepancies noted during the Apollo l3 mission.

14.1.1 Loss of Cryogenic Oxygen Tank 2 Pressure

At 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:


Question: What was the pressure at which the tank-2 relief-valve full-flow conditions occurred?

Answer: 1008 psia

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.


Question: How long did the pressure decrease last before the relief valve probably reseated?

Answer: 9 seconds

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.


Question: What was the immediate cause of the tank line to burst?

Answer: Burning of the wire insulation

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