Question: How long was data lost due to the panel separation shock?

Answer: 1.8 seconds

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


Question: What type of valve will be redesigned?

Answer: supply valve

supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems.

A more thorough discussion of this anomaly is presented in reference l.

14.1.2 Postlanding Vent Valve Malfunction

During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen.


Question: What type of valve was open during postlanding activities?

Answer: exhaust valve

supply valve will be redesigned to isolate polytetrafluoroethylenecoated wires from the oxygen. Warning systems at the Mission Control Center will be modified to provide more immediate and visible warnings of anomalies in all systems.

A more thorough discussion of this anomaly is presented in reference l.

14.1.2 Postlanding Vent Valve Malfunction

During postlanding activities, recovery personnel discovered that the postlanding ventilation inlet valve was closed and the exhaust valve wasopen.


Question: What was found to be within specifications in the valve-lock mechanism rigging?

Answer: Tolerances

The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle.

The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found.


Question: What happened when the malfunction was duplicated in the spacecraft?

Answer: Only partial travel of the handle

The valve-lock mechanism rigging tolerances were found to be within specifications. When reassembled in the spacecraft, the malfunction was duplicated with only partial travel of the handle.

The ventilation system was designed with two flexible control-cable assemblies linked to one handle, which is pulled to operate the two valves. An inherent characteristic of this design is that one control cable will nearly always slightly lag the other when the handle is pulled. At full extension of the handle, the travel in each cable assembly is more than sufficient to disengage both plungers and allow both valves to operate. Checkout procedures prior to flight were found to be satisfactory. There was no evidence of mechanical failure or malfunction nor were any outof-tolerance components found.


Question: What was the amplitude of the fluctuations observed in the computer readout of the optics shaft angle?

Answer: 0.3 degree

To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan.

14.l.3 Shaft Fluctuations in the Zero Optics Mode

Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission.


Question: What was the system mode when the fluctuations were observed?

Answer: zero optics mode

To guard against operational problems of this type in the future, a caution note has been added in the Apollo Operations Handbook to actuate the ventilation valve handle over its full travel before switching on the postlanding vent fan.

14.l.3 Shaft Fluctuations in the Zero Optics Mode

Beginning at approximately 4o hours, fluctuations of as much as 0.3 degree were observed in the computer readout of the optics shaft angle. The system had been powered up throughout the flight and had been in the zero optics mode since the star/horizon navigation sightings at 3l hours. Crew observation of the manual readout subsequently confirmed that the fluctuation was actually caused by motion of the shaft. The circumstances and time of occurrence were almost identical to a similar situation which occurred during the Apollo l2 mission.


Question: What type of resolver is used only for the shaft axis?

Answer: Half-speed resolver

The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation.

Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection.


Question: What is the equivalent of zero output in shaft rotation?

Answer: Zero degrees

The most likely candidate is the half-speed resolver, which is used only for the shaft axis and only to provide an unambiguous zero reference. The reference voltage is applied to the rotor through slip rings (fig. 14-3), connected as shown in figure 14-4. If any resistance is present in the common ground path through the slip ring, a portion of the reference voltage will appear across the quadrature winding and induce a finite output (different than zero). Zero output is equivalent to zero degrees in shaft rotation.

Simulated changes in slip ring impedance were made on the half-speed resolver in the shaft loop (fig. 14-4). An impedance of 50 ohms produced an offset of approximately plus 0.5 degree in sextant shaft angle. The trunnion loop does not use this type of resolver or connection.


Question: What was the angle in pitch that the Command Module Pilot had manually adjusted the antenna settings to?

Answer: 23 degrees

14.l.4 High-Gain Antenna Acquisition Problem

Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station.


Question: How far away from the line of sight to the ground station did the antenna boresight axis point due to the difference between the two sets of angles?

Answer: 35 degrees

14.l.4 High-Gain Antenna Acquisition Problem

Prior to the television transmission at approximately 55 hours, difficulty was experienced in obtaining high-gain antenna acquisition and tracking. The Command Module Pilot had manually adjusted the antenna settings to plus 23 degrees in pitch and 267 degrees in yaw, as requested by the ground 7 hours earlier. The most favorable settings for 55 hours were actually plus 5 degrees in pitch and 237 degrees in yaw. The difference between these two sets of angles pointed the antenna boresight axis approximately 35 degrees away from the line of sight to the ground station.


Question: What was the decrease in uplink signal strength when the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna?

Answer: 6 dB

When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10.


Question: In what mode was the high-gain antenna when the reacquisition mode was selected at 55:00:10?

Answer: medium-beam, manual mode

When the transmission was switched from the omnidirectional antenna to the manual mode of the high-gain antenna, there was a 6 dB decrease in uplink signal strength and a 17 dB decrease in dowmlink signal strength. With the high-gain antenna in the wide beam mode and nearly boresighted, the uplink and downlink signal strengths should have been at least equal to the signal strength obtained with an omnidirectional antenna. A comparison of the wide-, medium-, and narrow-beam transmit and receive pat-- terns indicates the high-gain antenna mode was in a medium-beam, manual mode at the time of acquisition and remained in this configuration until the reacquisition mode was selected at 55:00:10.


Question: What happened to the antenna when it acquired the earth in wide beam?

Answer: The antenna acquired the earth in wide beam.

limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual


Question: What triggered the switch to manual system?

Answer: the scan limit function line

limit function line just prior to point B, the antenna acquired the earth in wide beam. When the antenna reached the scan limit function line, the antenna control logic would switch the system to the manual


Question: What was the temperature at which the system was cold soaked for 7 hours?

Answer: 30° F

An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive.

A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests.

Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated.


Question: What was the result of the continuous functional tests while the system was warming up?

Answer: The system operated normally

An enlarged photograph of the scroll was examined in detail to determine if the scroll started properly. While no abnormal indications were observed, the interpretation of these data is not conclusive.

A complete functional test was performed and the flight problem could not be duplicated. The system was cold soaked for 7 hours at 30° F. While the system was slowly warming up, continuous functional tests were being performed to determine if thermal gradients could have caused the problem. The system operated normally throughout all tests.

Following verification of the light and sensing circuit, the mode switch was examined in detail. Tests were performed to determine contact resistance, and the switch was examined by X-ray for conductive contaminants and by dissection for nonconductive contaminants. No evidence of any switch problems was indicated.


Question: What was the temperature of the gas during operation?

Answer: $2000^{\circ}$ F

During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units.


Question: What was the pressure of the gas during operation?

Answer: 14 000 psi

During postflight inspection, it was discovered that propellant gas had leaked from the gusset-4 breech assembly, which is a part of the ape> cover jettison system (fig. 14-7).、 A hole was burned through the aluminum gusset cover plate (fig. l4-8), and the fiberglass pilot parachute mortar cover on the parachute side of the gusset was charred but not penetrated. The leakage occurred at the breech-plenum interface (fig. l4-9). The breech and plenum are bolted male and female parts which are sealed with a large O-ring backed up with a Teflon ring, as shown in figure l4-7. During operation, the breech pressure reaches approximately 14 000 psi and the gas temperature exceeds $2000^{\circ}$ F。The O-ring and backup ring were burned through and the metal parts were eroded by the hot gas at the leak path. The system is completely redundant in that either thruster system will effect apex cover jettison. No evidence of gas leakage existed on the previous firings of 56 units.


Question: What type of material will be applied to the interior of the breech plenum area on future spacecraft?

Answer: Polyimide

sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition.

Figure. l4-l0.- Tunnel gusset protection.

14.l.7 Reaction Control Isolation Valve Failure

During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation.


Question: Where did the miswiring of the fuel valve closing coil occur during initial installation?

Answer: On a terminal board

sheet (fig. l4-l0) will be applied to the interior of the breech plenum area on future spacecraft. The protection provided by the polyimide has been proof-tested by firing the assembly without the O-ring, simulating a worst-case condition.

Figure. l4-l0.- Tunnel gusset protection.

14.l.7 Reaction Control Isolation Valve Failure

During postflight decontamination of the command module reaction control system, the system l fuel isolation valve was found open when it should have been closed. All other propellant isolation valves were in the closed position. The subsequent failure investigation revealed that the lead from the fuel valve closing coil was wired to an unused pin on a terminal board instead of to the proper pin. X-rays of the terminal board and closeout photographs indicate the miswiring occurred during initial installation.


Question: What type of checks are used to verify the circuit in closeout installations?

Answer: Functional checks

Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired.

14.l.8 Potable Water Quantity Fluctuations


Question: What type of testing will be performed on future spacecraft to prove that the isolation valves are properly wired?

Answer: Resistance checks

Certain components are wired into the spacecraft wiring harness by inserting crimped, pinned ends of the wiring into terminal boards of the spacecraft harness. In many cases, this wiring is part of closeout installations and circuit verification can only be accomplished through functional checks of the component. This anomaly has pointed out the fact that circuits verified in this manner must be analyzed to determine if functional checks provide an adequate verification. All circuits have been analyzed with the result that the service module and command module reaction control system propellant isolation valves are the only components which require additional testing. Resistance checks will be performed on all future spacecraft to prove that the isolation valves are properly wired.

14.l.8 Potable Water Quantity Fluctuations


Question: What was the duration of the first water quantity fluctuation?

Answer: 5 minutes

The potable water quantity measurement fluctuated briefly on two occasions during the mission. At about 23 hours, the reading decreased from 98 to 79 percent for about 5 minutes and then returned to a normal reading of approximately l02 percent. Another fluctuation was noted at about 37 hours, at which time the reading decreased from its upper limit to 83.5 percent. The reading then returned to the upper limit in a period of 7 seconds.

Preflight fluctuations of from 2 to 6 percent near the full level were observed once during the countdown demonstration test, and a possible earlier fluctuation of about 4 percent at the half-full level was noted during the flight readiness test.


Question: At what time did the second water quantity fluctuation return to its normal reading?

Answer: in a period of 7 seconds

The potable water quantity measurement fluctuated briefly on two occasions during the mission. At about 23 hours, the reading decreased from 98 to 79 percent for about 5 minutes and then returned to a normal reading of approximately l02 percent. Another fluctuation was noted at about 37 hours, at which time the reading decreased from its upper limit to 83.5 percent. The reading then returned to the upper limit in a period of 7 seconds.

Preflight fluctuations of from 2 to 6 percent near the full level were observed once during the countdown demonstration test, and a possible earlier fluctuation of about 4 percent at the half-full level was noted during the flight readiness test.


Question: At what time did the suit pressure transducer reading suddenly drop from 6.7 to 5.7 psia?

Answer: 00:02:45

During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2).

(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure.


Question: What was the nominal regulated pressure of the cabin?

Answer: 5.0 psia

During launch the suit pressure transducer reading remained consistent with cabin pressure unti1 00:02:45, then suddenly dropped from 6.7 to 5.7 psia cQincidentally with S-II engine ignition (fig. 14-l2). The difference between the two measurements decreased to only 0.2 by l-l/2 hours, when the cabin reached its nominal regulated pressure of 5.0 psia. For this shirtsleeve mode, the suit and cabin pressure readings should be nearly equal. During normal variations in the command module cabin pressure, the suit pressure measurement responded sluggishly and indicated as much as l psi low. Subsequently, the measurement output decayed and remained in the 4.1 to 4.3 psia range for a cabin pressure of 5.0 psia until system deactivation at about 59 hours (fig. 14-l2).

(a)Lift-off through $4$ minutes. Figure l4-l2.- Suit and cabin pressure.


Question: What was noted to have escaped from the left-hand electrical circuit interrupter?

Answer: Propellant gas

During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3.


Question: What was found on the adjacent equipment due to the escaped propellant gas?

Answer: Soot

During postflight inspection of the command module, propellant gas was noted to have escaped from the left-hand electrical circuit interrupter, mounted in the lower equipment bay, and deposited soot on adjacent equipment. The right-hand circuit interrupter showed no evidence of a gas leakage. The removed breech, showing the displaced O-ring and crushed attenuator block, is shown in figure l4-l3.


Question: What is the percentage of the block that has been crushed when the O-ring enters the chamfer in the breech assembly?

Answer: 94

The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include:

a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block.


Question: What is the distance between the piston O-ring and the chamfer in the breech assembly when the fork contacts the attenuator block?

Answer: 0.075 inch

The worse-case tolerance buildup is when the fork contacts the attenuator block and the piston O-ring is 0.075 inch from entering the chamfer in the breech assembly. The O-ring enters this chamfer when the block has been crushed about 94 percent, at which point an O-ring displacement and accompanying gas escape could be expected. The factors which affect the degree of attenuator crushing are generally uncontrollable within narrow limits and include:

a. Sliding friction of the many electrical contact pins, the several camming and fork-to-plate surfaces, and the piston b. Forces exerted by the springs, which hold the lift and base plates together in the assembled position C. Propellant gas pressure and the simultaneous increase of pres sure in the two breeches and the plenum d. Simultaneous occurrence of the electrical firing signals to the two cartridges e. Physical properties of the attenuator block.


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

Answer: 7.9 psi/hour

helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively.

The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank.


Question: At what pressure range did the helium tank's rise-rate characteristics increase in the manner exhibited during the countdown demonstration test?

Answer: Between 640 and 900 psia

helium tanks, one at the manufacturer's plant and the other at the Manned Spacecraft Center. The results indicated pressure rise rates of 8.8 and 8.7 psi/hour, respectively.

The heat-leak test during the countdown demonstration indicated a normal rise rate of 7.9 psi/hour at 640 psia, whereas the special test showed an abnormal rise rate of 14.9 psi/hour above 900 psia. At some helium temperature equivalent to a pressure between 640 and 900 psia, the rise-rate characteristics would increase in the manner exhibited during the countdown demonstration test. Extrapolating these results to the flight conditions, it was determined that the helium tank was fully capable of supporting a lunar landing timeline, and the decision was made to proceed with the flight using the existing tank.


Question: What was the normal prelaunch-standby rise rate of the helium tank?

Answer: 7.8 psi/hour

The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard.

The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process.


Question: What material was used for the insulation between the two shells of the helium tank?

Answer: aluminized Mylar

The prelaunch-standby rise rate was a normal 7.8 psi/hour. During flight, the zero-g rise rate of 7 psi/hour was slightly higher than expected, but still satisfactory. Following the first descent engine firing at 6l-1/2 hours, the rise rate increased to 10.5 psi/hour, rather than returning to its normal value, as shown in figure 14-14.’ After the second firing at 79-l/2 hours, the rise rate again increased, this time to approximately 33 psi/hour until about 109 hours, when the helium-tank burst disc ruptured at l937 psia, as it should have and vented the remaining helium overboard.

The helium tank is a double-walled titanium pressure vessel, with 173 layers of aluminized Mylar insulation between the two shells. The annular region is evacuated to a level of $\mathsf{10}^{-\gamma}$ torr during the manufacturing process.


Question: What is the likely cause of the anomaly in the tank-insulation?

Answer: Tank-insulation degradation

The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during


Question: What happens to the contaminant when it freezes?

Answer: The pressure of the contaminant is too low to significantly affect the thermal conductivity.

The most likely cause of the anomaly is a tank-insulation degradation which would result in increased heat conduction to the helium. The insulating characteristics of the annular vacuum in tank was most likely degraded by the introduction of a contaminant (probably hydrogen) in extremely small concentrations (approximately $10^{-}{\bar{6}}$ pounds). These contaminants when vaporized can exponentially increase the thermal conductivity in proportion to their vapor pressure, as indicated by special tests. While loading helium into the tank, the contaminants wouid freeze upon the inner shell.. In the frozen state, the pressure of the contaminant is too low to significantly affect the thermal conductivity. However, the flow check which preceded the cold-soak operation would vaporize the contaminants in the vicinity of the heat exchanger lines which pass through the annulus. The subsequent increase in thermal conductivity could cause the abnormally high pressure-rise rate observed during


Question: What is the desired steady-state pressure rise rate for a perfect tank over the entire range of temperatures?

Answer: 8 psi/hour

Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure.

A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity.


Question: How long after pumpdown is the pressure in the jacket measured to verify vacuum integrity?

Answer: 2 or 3 weeks

Figure $_{14-14}$ .- Inflight profile of supercritical helium tank pressure.

A screening test was devised for all future flight tanks to supplement normal helium tank testing. The purpose of this test is to determine the pressure rise rate for a wide range of helium temperatures from approximately $9^{\circ}$ to $\bar{\mathsf{1}}\bar{2}\bar{3}^{\circ}$ R. .For a perfect tank, the steady-state.rise rate should remain at approximately 8 psi/hour over the entire range of temperatures. The Apollo 14, l5, and l6 tanks have been subjected to the screening test, and each exhibit the same phenomena observed during Apollo l3, but to a lesser degree. For new tanks, the manufacturer will periodically analyze the gases removed from the vacuum jacket during pump down for possible contaminants. The pressure in the jacket will be measured 2 or 3 weeks after pumpdown to verify vacuum integrity.


Question: What was the duration of the current transients in the descent batteries?

Answer: 2 seconds

At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down.

Figure 14-l5.- Descent stage battery location.

The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition.


Question: Where was the momentary short circuit located in the dc electrical system?

Answer: Lunar-Module-Pilot side

At 97 hours 14 minutes, the crew reported a thumping noise and snowflakes venting from quadrant 4 of the lunar module descent stage (fig. l4-l5). All four descent batteries experienced current transients at 97:13:53 for about 2 seconds, with corresponding drops in dc bus voltage (fig. l4-l6). Also, the water glycol pressure differential for the heat transport system decreased momentarily, indicating that the glycol pump momentarily slowed down.

Figure 14-l5.- Descent stage battery location.

The thumping noise occurred at about the same time as the current spikes._ The current spikes show that a momentary short circuit existed in the Lunar-Module-Pilot side of the dc electrical system, which includes descent batteries l and 2 (fig. 14-l6). The current surge was not of sufficient duration either to open the balance-load cross-tie circuit breakers, to display a reverse current indication, or to trip a batteryoff relay as a result of an overcurrent condition.


Question: What percentage of the total current load was battery 2 providing immediately after the current surges?

Answer: 80 percent

The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section.


Question: At what time did battery 2 give an indication of a battery malfunction?

Answer: 99:5l:09

The data show that descent battery 2 experienced at least a 60-ampere current surge. This condition could have been a reverse current into the battery, since the instrumentation system does not indicate the direction of current. Immediately after the current surges, battery l current returned to its original value while battery 2 provided about 80 percent of the total current load. After sustaining a surge load, the battery terminal voltage normally increases for a short period of time. Since battery 2 experienced the highest surge, it should have temporarily assumed the most load. Within l0 minutes all batteries were properly sharing the current load, and no subsequent abnormal performance was observed. At 99:5l:09, battery 2 gave an indication of a battery malfunction, discussed in more detail in the next section.


Question: What was the maximum differential pressure that caused the leak in the oxygen tank 2?

Answer: 193 psi

During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction.

The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring


Question: How much leakage was allowed for the valve in either direction?

Answer: 360 scc/hr

During the flight, the pressure in the ascent stage oxygen tank 2 increased, indicating a reverse leakage through the shutoff valve from the oxygen manifold (fig. 14-l9) into the tank. The leak rate, with a maximum differential pressure of 193 psi, varied from about 0.22 lb/hr (70 0oo scc/hr) to zero when the tank pressure reached manifold pressure. Allowable leakage for the valve in either direction is 360 scc/hr. Preflight test data indicate a reverse leakage of 360 scc/hr and no excessive leaking in the forward direction.

The internal portion of three valves of this type had been replaced previously on the spacecraft because of excessive leakage through the ascent oxygen tank l shutoff valve. In one valve, a roiled O-ring


Question: What is the recommended compression range for the O-ring by the manufacturer?

Answer: between 0.01l5 and 0.0225 inch

(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination.

The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly.

Figure l4-20.- Ascent stage tank shutoff valve:.


Question: What is the cause of the leakage in the other two valves?

Answer: contamination

(fig. 14-20) caused the leakage. When the valve is installed, the forward O-ring can be rolled and damaged when it passes the manifold port. In the other two valves, the cause was not identified and was assumed to be contamination.

The production tolerances of the valve and bore were examined to determine if a tolerance buildup problem existed. The manufacturer's specification to which the valve was designed requires that the O-ring be subjected to a compression of between 0.01l5 and 0.0225 inch, whereas the 0-ring supplier recommends between 0.0ll and 0.017 inch. The added compression allowed in the valve design would aggravate the tendancy for the O-ring to roll during valve .assembly.

Figure l4-20.- Ascent stage tank shutoff valve:.


Question: What type of material is stitched to the inner surface of the window shade?

Answer: A Beta Cloth

Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves.

The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade.

Figure 14-21.- Cracked left-hand window shade.

Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight.


Question: What percentage of stitch holes had cracks extending from them on the window shade?

Answer: 80 percent

Leak tests previously performed on the valve were inadequate, in that only reverse leakage at high pressure was determined. For future vehicles, forward and reverse leakage at both high and low pressures will be measured to detect any defective valves.

The left-hand window shade showed three large separations when it was first placed in the stowed position during flight (fig. l4-2l). A Beta Cloth backing is stitched to the inner surface of the. Aclar shade. The cracks propagated from the sewing stitch holes on the periphery of the shade. About 1/8-inch-long cracks extended from about 80 percent of the stitch holes in a direction parallel with the curl axis of the shade.

Figure 14-21.- Cracked left-hand window shade.

Cracking as a result of Aclar embrittlement has occurred before, therefore, the Apollo l3 shades were examined prior to flight. Since no cracks were found, the shades were approved for flight.


Question: What percentage of elongation does the modified Aclar material provide?

Answer: 25 percent

The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight.

14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera

For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies.


Question: What type of material will the shades for future vehicles be fabricated from?

Answer: Aclar

The Aclar supplier has developed a heating and quenching process to provide material with an elongation in excess of 25 percent, as compared to elongations of from 6 to l2 percent for the failed shades. Shades for future vehicles will be fabricated from this more ductile material. The Aclar will be reinforced with Mylar tape before the Beta Cloth backing is stitched to the shade. The modified shades have been requalified for the next flight.

14.3.l Loose Lens Bumper On Lunar Module 16-mm Camera

For launch, the l6-mm camera is mounted to point through the Lunar Module Pilot's window with the l0-mm lens and bumper attached. At the time of inflight lunar module inspection, the bumper was found to have separated from the camera lens. The bumper was replaced and remained attached for the remainder of the flight. Looseness has been experienced during previous lens /bumper assemblies.


Question: What was used to secure the set screw of the interval timer set knob prior to the change?

Answer: A special gripping compound

To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly.

14.3.2 Failure of the Interval Timer Set Knob

The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin.


Question: What will be used to secure the knobs on timers for future flights?

Answer: A roll pin

To prevent recurrence of the problem, the mating surface of the bumper will be swaged for future missions so as to provide an interference fit with the internal surface threads of the l0-mm lens assembly.

14.3.2 Failure of the Interval Timer Set Knob

The onboard interval timer, which has two timing ranges (o to 6 and 0 to 60 minutes), is stowed in the command module for crew use in timing such routine functions as fuel cell purges, cryogenic system fan cycles, and so forth. A tone advises the crew when the set time period has elapsed. Prior to 55 hours, the time-period set knob came off in a crewman's hand because of a loosened set screw. The set screw had been secured with a special gripping compound. Postflight examination of other flight timers indicated that this compound apparently does not provide a strong enough retention force for this application. Therefore, the knobs on timers for future flights will be secured to the shaft with a roll pin.


Question: What type of drops will be packaged the same as eye drops in medical kits for future flights?

Answer: Nose drops

Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops.

The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report.


Question: What was the Apollo mission that required an emergency abort?

Answer: Apollo l3

Medical kits for future flights will include nose drops packaged the same as the eye drops. This packaging has been satisfactory on previous flight for eye drops.

The Apollo l3 mission was the first in the Program requiring an emergency abort, with the Gemini VIlI mission the only prior case in manned spaceflight where a flight was terminated early. The excellent performance of the lunar module systems in a backup capacity and the training of both the flight crew and ground support personnel resulted in the safe and efficient return of the crew. The following conclusions are drawn from the information contained in this report.


Question: What was the purpose of the lunar flyby mission?

Answer: a lunar flyby mission, including three planned experiments

d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency.

e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived.

The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented.


Question: What was the outcome of the mission regarding the lunar module's backup capability?

Answer: information which would have otherwise been unavailable

d. The effectiveness of preflight crew training, especially in conjunction with ground personnel, was reflected in the skili and precision with which the crew responded to the emergency.

e. Although the mission was not a complete success, a lunar flyby mission, including three planned experiments (lightning phenomena, earth photography, and S-IVB lunar impact), was completed and information which would have otherwise been unavailable, regarding the long-term backup capability of the lunar module, was derived.

The configuration of the Apollo l3 spacecraft and launch vehicle was nearly identical to that of Apollo l2, and the spacecraft/launch vehicle adapter and launch escape system underwent no changes. The few changes to the command and service modules and the lunar module are discussed in the following paragraphs. A discussion of the changes to the Apollo lunar surface experiments package and a listing of the spacecraft mass properties are also presented.


Question: What was the original location of the sensing point for the water separator drain tank?

Answer: The carbon dioxide sensor

propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush


Question: What was the purpose of adding a removable flow limiter to the inlet for the primary lithium hydroxide cartridge?

Answer: To reduce the water separator speed and to minimize the possibility of condensed water in the suit.

propulsion system was changed to include a bypass line around the fuel/helium heat exchanger such that if the heat exchanger should freeze during venting, pressures would equalize on both sides of the heat exchanger. The sensing point for the water separator drain tank was changed from the location of the carbon dioxide sensor to a point upstream of the suit fans, thus eliminating migration of water to the carbon dioxide sensor and improving its operation. A removable flow limiter was added to the inlet for the primary lithium hydroxide cartridge to reduce the water separator speed and to minimize the possibility of condensed water in the suit. A dust filter was incorporated at the inlet of the cabin fan to reduce the amount of free lunar dust in the cabin. Redesigned water/ glycol and oxygen disconnects having redundant seals were installed to improve reliability and to permit up to 5 degrees of connector misalignment. To decrease the possibility of lunar dust contamination, a brush


Question: What material were the bristles on the vacuum brush changed from?

Answer: Teflon

was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag.


Question: What was added to the lunar sample tote bag?

Answer: a cover

was added for cleaning the suits before ingress, the bristles on the vacuum brush were changed from Teflon to Nylon, and a cover was added to the lunar sample tote bag.


Question: What instruments were deleted from the Apollo 13 package compared to Apollo 12?

Answer: Solar wind spectrometer, magnetometer, and suprathermal ion detector

The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13.

The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay.

Figure A-l.- Experiment subpackage number l.

The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements.


Question: What was the purpose of the heat flow experiment on Apollo 13?

Answer: To measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials

The Apollo lunar surface experiment package stowed for Apollo l3 was similar to that for Apollo l2. However, the solar wind spectrometer, magnetometer, and suprathermal ion detector, included on Apollo l2, were deleted from Apollo 13. A heat flow experiment and a charged particle environment detector were added for Apollo l3. The cold-cathode ion gage experiment deployed during Apollo l2 was significantly modified for Apo1lo13.

The Apollo lunar surface experiments package consisted of two subpackages as shown in figures A-l and A-2. These were stowed in the lunar module scientific equipment bay.

Figure A-l.- Experiment subpackage number l.

The heat flow experiment was designed to measure the thermal gradient of the upper 3 meters of the lunar crust and the thermal conductivity of the lunar surface materials. Lunar heat flow calculations could be based on the measurements.


Question: What type of insulation was used in the S-Il stage?

Answer: Spray foam

The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover.

Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation.


Question: Where were telemetry measurements relocated to provide a more complete analysis of platform vibrations?

Answer: In the inertial platform

The experiment consisted of an electronics package with sunshield and reflector, to shade the thermal plate from the direct sunlight, and a sensor package with aperture and dust cover.

Spray foam was used exclusively as insulation in the S-Il stage to reduce weight. A fourth battery was installed in the instrument unit to extend the tracking capability to lunar distance in support of the S-IVB lunar impact experiment. Telemetry measurements in the inertial platform were added and, in some cases, were relocated to provide a more complete analysis of platform vibrations. Four wires were added to the distributor in the emergency detection system, located in the instrument unit, to provide automatic ground command capability at spacecraft separation in the event of a contingency separation.


Question: What is the moment of inertia of the Command&service modules?

Answer: 847.4


Question: What is the product of inertia of the Lunar module before the cryotenic oxygen tank ignition?

Answer: 33995


Question: What was the ignition cutoff for the third midcourse correction?

Answer: 398.7


Question: What was the altitude of the command module after lunar module separation?

Answer: 2362


Question: What is the first number in the list?

Answer: 1038.6


Question: What is the last word in the list?

Answer: Landing


Question: Where was the history of command and service module (cSM 109) operations at the manufacturer's facility shown?

Answer: Downey, California

ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases.

The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2.

The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4.

Figure B-l.- Checkout flow for command and service modules at contractor's facility.

Figure B-2.- Command and service module checkout history at Kennedy Space Center.

Figure B-3.- Checkout flow for lunar module at contractor's facility.

Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center.


Question: Where was the history of the lunar module (LM-7) operations at Kennedy Space Center shown?

Answer: Figure B-4

ALunar module was docked to the command module from initial docking wntil just prior to entry. "Mass properties are referenced to the coordinate system of the lnar module, which provided spacecraft dynanic control during these phases.

The history of command and service module (cSM 109) operations at the manufacturer's facility, Downey, California, is shown in figure B-l, and the operations at Kennedy Space Center, Florida, in figure B-2.

The history of the lumar module (LM-7) at the manufacturer's facility, Bethpage, New York, is shown in figure B-3, and the operations at Kennedy Space Center, Florida, in figure B-4.

Figure B-l.- Checkout flow for command and service modules at contractor's facility.

Figure B-2.- Command and service module checkout history at Kennedy Space Center.

Figure B-3.- Checkout flow for lunar module at contractor's facility.

Figure $\mathbb{R}{-}\mathbb{4}$ .- Lumar module checkout history at Kennedy Space Center.


Question: What is the time range of the first recorded data point?

Answer: 57:57


Question: What is the type of data recorded in the last entry of the table?

Answer: MSFN


Question: What was the publication date of the report titled "Reaction Control System Performance"?

Answer: August 1969

Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command


Question: How many reports were published for the Apollo 8 mission?

Answer: 7

Supplement number Title Publi cation date/status Apollo 7 1 2 Trajectory Reconstruction and Analysis Communication System Performance May 1969 June 1969 3 Guidance, Navigation, and Control System Performance Analysis November 1969 4 5 Reaction Control System Performance Cancelled August 1969 6 Entry Postflight Analysis December 1969 Apollo 8 1 Trajectory Reconstruction and Analysis December 1969 2 Guidance, Navigation, and Control System Performance Analysis November 1969 3 Performance of Command and Service Module March 1970 4 Reaction Control System Service Propulsion System Final Flight September 1970 6 Evaluation Analysis of Apollo 8 Photography and December 1969 7 Visual Observations Entry Postflight Analysis December 1969 Apollo 9 1 Trajectory Reconstruction and Analysis November 1969 2 Command and Service Module Guidance, Navi- gation, and Control System Performance November 1969 3 Lunar Module Abort Guidance System Perform- ance Analysis November 1969 4 Performance of Command


Question: What was the year the Ascent Propulsion System received its Final Flight Evaluation?

Answer: 1969

and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969


Question: What was the status of the item listed as number 9?

Answer: Cancelled

and Service Module Reaction Control System Apri1 1970 5 Service Propulsion System Final Flight Evaluation December 1969 6 Performance of Lunar Module Reaction Control System Final review 7 Ascent Propulsion System Final Flight Evaluation December 1969 8 Descent Propulsion System Final Flight Evaluation September 1970 9 Cancelled 10 Stroking Test Analysis December 1969 11 Communications System Performance December 1969 12 Entry Postflight Analysis December 1969


Question: What was the publication date of the report titled "Performance Analysis"?

Answer: December 1969

Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review


Question: What was the status of the report titled "Analysis of Apollo lo Photography and Visual Observations"?

Answer: In publication

Supplement number Title Publication date/status Apollo 10 1 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System March 1970 2 Performance Analysis December 1969 3 Perfornance of Command and Service Module Reaction Control System Final review 7 Service Propulsion System Final Flight September 1970 5 Evaluati on Performance of Lunar Module Reaction Control Final review 6 System Ascent Propulsion System Final Flight January 1970 7 Evaluation Descent Propulsion System Final Flight January 1970 8 Evaluati on Cancelled 9 Analysis of Apollo lo Photography and Visual Observations In publication 10 11 Entry Postflight Analysis Communi cations System Performance December 1969 December 1969 Apollo ）11 2 3 4 Trajectory Reconstruction and Analysis Guidance, Navigation, and Control System Performance Analysis Performance of Command and Service Module Reaction Control System Service Propulsion System Final Flight Evaluation May 1970 September 1970 Review Review


Question: What was the title of the report about the Apollo 12 trajectory?

Answer: Apollo 12 Trajectory Reconstruction and Analysis

Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation

Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970.

Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970.


Question: What was the status of the report "Apollo 12 Trajectory Reconstruction and Analysis" when it was published?

Answer: 1

Supplement number Title Publication date/status Apollo 12 Trajectory Reconstruction and Analysis 1 2 3 Guidance, Navigation, and Control System Performance Analysis Service Propulsion System Final Flight September 1970 September 1970 4 Evaluation Ascent Propulsion System Final Flight Evaluation Preparation Preparation 5 6 Descent Propulsion System Final Flight Evaluation Preparation 7 Apollo l2 Preliminary Science Report Landing Site Selection Processes July 1970 Final review Apollo 13 1 Guidance, Navigation, and Control System Performance Analysis Review 2 Descent Propulsion System Final Flight Evaluation Entry Postflight Analysis Preparation

Manned Spacecraft Center: Apollo 13 Cryogenic Oxygen Tank 2 Anomaly Report. MSC-02545. June 1970.

Marshall Space Flight Center: Saturn V Launch Vehicle Flight Evaluation Report AS-508 Apollo 13 Mission. MPR-SAT-FE-70-2. June 1970.


Question: What year was the Study Conference on the Global Atmospheric Research Program held?

Answer: 1967

Marshall Space Flight Center, Kennedy Space Center, Manned Spacecraft Center: Analysis of Apollo l2 Lightning Incident, MSC-01540. February 1970.

ICSU/IUGG Committee on Atmospheric Sciences: Report of the Study Conference on the Global Atmospheric Research Program, 1967.

Bulletin of the American Meteorological Society, Vol. 50, No. 7: Cloud Height Contouring from Apollo 6 Photography, by V. S. Whitehead, I. D. Browne, and J. G. Garcia. 1969.

Defense Supply Agency, Washington, D. C.: Military Standardization Handbook_ Optical Design, MIL HDBK-14l. 1962.

NASA Headquarters: Apollo Flight Mission Assignments. OMSF M-D MA500-11 (SE 010-000-1). 0ctober 1969.

Manned Spacecraft Center: Mission Requirement, H-2 Type Mission (Lunar Landing). SPD9-R-053. November 10, 1969.


Question: Who authored the article "Cloud Height Contouring from Apollo 6 Photography"?

Answer: V. S. Whitehead, I. D. Browne, and J. G. Garcia

Marshall Space Flight Center, Kennedy Space Center, Manned Spacecraft Center: Analysis of Apollo l2 Lightning Incident, MSC-01540. February 1970.

ICSU/IUGG Committee on Atmospheric Sciences: Report of the Study Conference on the Global Atmospheric Research Program, 1967.

Bulletin of the American Meteorological Society, Vol. 50, No. 7: Cloud Height Contouring from Apollo 6 Photography, by V. S. Whitehead, I. D. Browne, and J. G. Garcia. 1969.

Defense Supply Agency, Washington, D. C.: Military Standardization Handbook_ Optical Design, MIL HDBK-14l. 1962.

NASA Headquarters: Apollo Flight Mission Assignments. OMSF M-D MA500-11 (SE 010-000-1). 0ctober 1969.

Manned Spacecraft Center: Mission Requirement, H-2 Type Mission (Lunar Landing). SPD9-R-053. November 10, 1969.


Question: What was the launch date of the Apollo 4 mission?

Answer: Nov. 9, 1967

Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18


Question: What was the purpose of the Apollo 6 mission?

Answer: Verification of closed-loop emergency detection system

Mi ssion. Apollo 4 Spacecraft Description Launch date Launch site SC-017 LTA-10R Supercircular entry at lunar. return velocity Nov. 9, 1967 Kennedy Space Center, Fla. Apollo 5 LM-1 First lunar module flight Jan. 22, 1968 Cape Kennedy, Fl&. Apollo 6 SC-020 LTA-2R Verification of closed-loop emergency detection system Apri1 4, 1968 Kennedy Space Center, Fia. Apo1lo7 CSM 101 First manned flight; earth-orbital 0ct.11，1968 Cape Kennedy, Fl&. Apollo 8 CSM 103 First manned lumar orbital flight; first manned Saturn V launch Dec.21,1968 Kennedy Space Apollo 9 CSM 104 LM-3 First manned lunar module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center,Fla. Apollo 10 CSM 106 t-WT First lumar orbit rendezvoua; low pass over lunar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 LM-6 Second lunar landing Nov. 14, 1969 Center, Fla. Kennedy Space Center. F18


Question: What is the location of the NASA-Manned Spacecraft Center?

Answer: Houston, Texas 77058

POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION

NASA-Manned Spacecraft Center Houston, Texas 77058


Question: What is the recipient of the letter?

Answer: PT2

POSTAGE AND FEES PAID NATIONAL AERONAUTICS AND SACE ADMINISTRATION

NASA-Manned Spacecraft Center Houston, Texas 77058


Question: What was the date of the first lunar landing?

Answer: July 16, 1969

Mi ssion Spacecraft Description Launch date Launch site Apollo4 SC-017 LTA-10R Supercircular entry at lunar Nov.9,1967 Kennedy Space Center, Fla. Apollo 5 LM-1 return velocity First lunar module flight Jan.22,1968 Cape Kennedy, Apollo 6 SC-020 LTA-2R Verification of closed-loop April 4, 1968 Fla. Kennedy Space Center, Fla. Apollo7 CSM 101 emergency detection system First manned flight; Oct.11，1968 Apol1o 8 CSM 103 earth-orbital First manned lunar Dec.2l,1968 Cape Kennedy, Fla. Kennedy Space Apol1o9 CSM 104 orbital flight; first manned Saturn V launch First manned lunar Apollo 10 LM-3 module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center, Fla. CSM 106 LM-4 First lunar orbit rendezvous; low pass over lumar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 Second lunar landing Center, Fla. Apollo 13 LM-6 Nov. 14, 1969 Kennedy Space Center, Fla.


Question: What was the name of the launch site for Apollo 6?

Answer: Kennedy Space Center, Fla.

Mi ssion Spacecraft Description Launch date Launch site Apollo4 SC-017 LTA-10R Supercircular entry at lunar Nov.9,1967 Kennedy Space Center, Fla. Apollo 5 LM-1 return velocity First lunar module flight Jan.22,1968 Cape Kennedy, Apollo 6 SC-020 LTA-2R Verification of closed-loop April 4, 1968 Fla. Kennedy Space Center, Fla. Apollo7 CSM 101 emergency detection system First manned flight; Oct.11，1968 Apol1o 8 CSM 103 earth-orbital First manned lunar Dec.2l,1968 Cape Kennedy, Fla. Kennedy Space Apol1o9 CSM 104 orbital flight; first manned Saturn V launch First manned lunar Apollo 10 LM-3 module flight; earth orbit rendezvous; EVA Mar.3,1969 Kennedy Space Center, Fla. CSM 106 LM-4 First lunar orbit rendezvous; low pass over lumar surface May 18, 1969 Kennedy Space Center, Fla. Apollo 11 CSM 107 LM-5 First lunar landing July 16, 1969 Kennedy Space Apollo 12 CSM 108 Second lunar landing Center, Fla. Apollo 13 LM-6 Nov. 14, 1969 Kennedy Space Center, Fla.