8.0 LUNAR MODULE PERFORMANCE

8.1 STRUCTURAL AND MECHANICAL SYSTEMS

Lunar module structural loads were within design values for all phases of the mission. The structural assessment was based on guidance and control data, cabin pressure measurements, command module acceleration data, photographs, and crew comments.

Based on measured command module accelerations and on simulations using actual launch wind data, lunar module loads were determined to be within structural limits during earth launch and translunar injection. The sequence film from the onboard camera showed no evidence of structural oscillations during lunar touchdown, and crew comments agree with this assessment.

Landing on the lunar surface occurred with estimated landing velocities of 3.1 ft/sec vertical, 1.7 ft/sec in the plus-Y footpad direction, and 1.7 ft/sec in the plus-Z footpad direction. The spacecraft rates and attitude at touchdown are shown in figure 8-1. The minus-Y footpad apparently touched first, followed by the minus-Z footpad approximately 0.4 second later. The plus-Y and plus-Z footpads followed within 2 seconds and the vehicle came to rest with attitudes of 1.8 degrees pitch down, 6.9 degrees roll to the right and 1.4 degrees yaw to the left of west. Very little, if any, of the vehicle attitude was due to landing gear stroking. The final rest attitude of approximately 7 degrees was due almost entirely to local undulations at the landing point (fig. 8-2). From a time history of the descent engine chamber pressure, it appears that descent engine shutdown was initiated after first footpad contact but before plus-Y footpad contact. The chamber pressure was in a state of decay at 108:15:11, and all vehicle motion had ceased 1.6 seconds later.

Figure 8-1 - Attitude errors and rates during lunar landing sequence

Figure 8-2 - Lunar module landing site


Flight data from the guidance and propulsion systems were used in performing engineering simulations of the touchdown phase. As in Apollo 11 and Apollo 12, these simulations and photographs indicate that landing gear stroking was minimal if it occurred at all. Photographs also indicate no significant damage to the landing gear thermal insulation.

Sixteen-millimeter films taken from the command module prior to lunar-orbit docking support a visual observation by the crew that a strip of material about 4 feet long was hanging from the ascent stage base heat shield area. The base heat shield are~a is designed to protect the ascent stage from the pressure and thermal environment resulting from ascent engine plume impingement during staging. The absence of abnormal thermal responses in the ascent stage indicates that the heat shield was fully effective. Similar conditions have occurred during qualification tests whereby one or more layers of the heat shield material have become unattached. In these cases, the thermal effectiveness of the heat shield was not reduced.

8.2 ELECTRICAL POWER

The electrical power distribution system and battery performance was satisfactory with one exception, the ascent battery 5 open-circuit voltage decayed from 37.0 volts at launch to 36.7 volts at housekeeping, but with no effect on operational performance. All power switchovers were accomplished as requiredi and parallel operation of the descent and ascent batteries was within acceptable limits. The dc bus voltage was maintained above 29.0 volts, and maximum observed current was 73 amperes during powered descent initiation.

The battery energy usage throughout the lunar module flight is given in section 8.11.6. The ascent battery 5 open-circuit low voltage is discussed in section 14.2.1.

8.3 COMUNICATIONS EQUIPMENT

S-band steerable antenna operation prior to lunar landing was intermittent. Although antenna operation during revolution 13 was nominal, acquisition and/or tracking problems were experienced during revolutions 11 and 12. Acquisition was attempted but a signal was not acquired during the first 3 minutes after ground acquisition of signal on revolution 14. Because of this, the omnidirectional antennas were used for lunar landing. The steerable antenna was used for the ascent and rendezvous phase and the antenna performed normally. The problems with the steerable antenna are discussed in section 14.2.3.

Prior to the first extravehicular period, difficulty was experienced when configuring the communication system for extravehicular activity because of an open audio-center circuit breaker. Extravehicular communications were normal after the circuit breaker was closed.

During the latter part of the first extravehicular period, the television resolution decreased. The symptoms of the problem were indicative of an overheated focus coil current regulator. This condition, while not causing a complete failure of the camera, resulted in defocusing of the electron readout beam in the television tube and, consequently, a degradation of resolution. The high-temperature condition was caused by operating the camera for about 1 hour and 20 minutes while it was within the thermal environment of the closed modular equipment stowage assembly. The camera was turned off between the extravehicular periods to allow cooling. Picture resolution during the second extravehicular activity was satisfactory.

The VHF system performance was poor from prior to lunar liftoff through terminal phase initiation. This problem is discussed in detail in sections 7.4 and 14.1.4.

8. 4 RADAR

The landing radar self-test was performed at 105 hours 44 minutes, and the radar was turned on for the powered descent about 2 hours later. Four minutes fifty seconds prior to powered descent initiation, the radar changed from high- to low-scale. At that time, the orbital altitude of the lunar module was about 10.9 miles (referenced to landing site elevation). This condition prevented acquisition of ranging signals at slant ranges greater than 3500 feet, and velocity signals at altitudes greater than about 4600 feet. The radar was returned to high-scale by recycling the circuit breaker. A detailed discussion of this problem is given in section 14.2.4. Range and velocity performance from a slant range of about 25 000 feet to touchdown is shown in figure 14-22. There were no zero Doppler dropouts and no evidence of radar lockup resulting from particles scattered by the engine exhaust plume during lunar landing.

Rendezvous radar performance was nominal in all respects, including self-tests, checkout, rendezvous and lunar surface tracking, and temperature.

8.5 INSTRUMENTATION

The instrumentation system performed normally throughout the flight with the exception of three of the four ascent helium tank pressure measurements (two primary and two redundant). Coincident with propulsion system pressurization, these measurements exhibited negative shifts of up to 4 percent. The largest shifts were in the redundant measurements. These transducer shifts were caused by the shock induced by the pyrotechnically operated isolation valves. Since these measurements are used to monitor for leaks prior to propulsion system pressurization, a shift in these measurements at the time of system pressurization will not affect future missions. (See appendix A, section A.2-3, for a description of changes made subsequent to Apollo 13.)

8.6 GUIDANCE, NAVIGATION, AND CONTROL

At approximately 102 hours, the primary guidance system was turned on, the computer digital clock was initialized, and the platform was aligned to the command module platform. Table 8-I is a summary of the primary guidance platform alignment data. The abort guidance system was turned on at 102 hours 40 minutes and the attitude reference aligned to the lunar module platform. Table 8-II is a summary of inertial measurement unit component errors measured prior to launch and in flight. The abort guidance system was aligned to the primary guidance system six times, but data were available for only five, and are shown in table 8-III. Also shown in table 8-III are data from the independent alignment of the abort system performed in preparation for lunar lift-off. The abort guidance system had been aligned to the gravity vector and an azimuth angle supplied by the ground. Twenty-seven minutes later, just before lift-off, the abort system compared well with the primary system which had been inertially aligned to the predicted local vertical orientation for lift-off.

TABLE 8-I - LUNAR MODULE PLATFORM ALIGNMENT SUMMARY

Time,
hr:min
Type
alignment
Alignment mode Telescope
detent#/
star used##
Star angle
difference,
deg
Gyro torquing angle, deg Gyro drift, meru
Option* Technique** X Y Z X Y Z
102:58 - Docked alignment - - 0.009 0.029 -0.052 -0.5 -1.5 -2.8
105:09 P52 3 NA 2/22; 2/16 0.04 0.030 -0.038 0.028 - - -
105:27 P52 3 NA -- -- - 0.097 0.062 0.013 -1.5 2.0 -0.6
109:17 P57 3 1 NA NA 0.03 -0.016 0.015 -0.113 - - -
109:46 P57 3 2 -2/31; 6/00 0.02 -0.041 0.003 -0.054 1.0 -0.1 -1.4
110:05 P57 3 2 2/26; 6/00 -0.07 0.018 0.047 -0.121 - - -
129:56 P57 4 3 -- -- 0.01 0.044 0.069 -0.460 - - -
141:53 P57 4 3 -- -- 0.02 0.119 0.135 -0.349 -0.7 -0.8 -1.9
* 1 - Preferred; 2 - Nominal; 3 - REFSMMAT; 4 - Landing site

** 0 - Anytime; 1 - REFSMMAT plus g; 2 - Two bodies; 3 - One body plus g

# 1 - Left front; 2 - Front; 3 - Right front; 4 - Right rear; 5 - Rear; 6 - Left rear

## Star names: 00 Pollux; 16 Procyon; 22 Regulus; 26 Spica; 31 Arcturus




TABLE 8-IIA - INERTIAL COMPONENT HISTORY - LUNAR MODULE ACCELEROMETERS

Error Sample
mean
Standard
deviation
Number
of
samples
Countdown
value
Flight
load
Inflight performance
Power-up
to landing
Surface
power-up
to lift-off
Lift-off to
rendezvous
X - Scale factor error, ppm

Bias, cm/sec/sec
-895

1.27
36

0.05
6

6
-922

1. 26
-950

1.30
-

1.27
-

1.38
-

1.36
Y - Scale factor error, ppm

Bias, cm/sec/sec
-1678

1.63
79

0.03
9

9
-1772

1.61
-1860

1.65
-

1.62
-

1.74
-

1.71
Z - Scale factor error, ppm

Bias, cm/sec/sec
-637

1.39
25

0.02
6

6
-643

1.41
-670

1.39
-

1.35
-

1.46
-

1.45


TABLE 8-IIB - INERTIAL COMPONENT HISTORY - LUNAR MODULE GYROSCOPES

Error Sample
mean
Standard
deviation
Number
of
samples
Countdown
value
Flight
load
Inflight
performance
X Null bias drift, meru

Acceleration drift,
spin reference axis,
meru/g

Acceleration drift,
input axis,
meru/g
0.8



0.2



4.0
0.4



0.8



2.8
6



6



6
0.0



1.1



2.9
0.9



0



3.0
-1.9



-



-
Y Null bias drift, meru

Acceleration drift,
spin reference axis,
meru/g

Acceleration drift,
input axis,
meru/g
-2.8



3.0



-9.6
0.6



1.3



4.0
6



6



12
-3.6



4.5



-7.5
-2.7



3.0



-12.0
0.3



-



-
Z Null bias drift, meru

Acceleration drift,
spin reference axis,
meru/g

Acceleration drift,
input axis,
meru/g
-1.1



4.5



5.8
0.9



0.4



1.4
6



6



6
-1.1



4.5



7.2
-0.3



5.0



6.0
-0.5



-



-


TABLE 8-III - GUIDANCE SYSTEMS ALIGNMENT COMPARISON

Time
of
alignment
Primary minus abort system
Alignment error (degrees)
X Y Z
103:54:44.99 0.000 0.003 0.014
104:04:45.9 0.061 0.030 0.002
104:34:45.2 0.000 0.007 0.003
109:28:36 -0.002 0.034 0.000
141:15:25.2 0.000 0.002 0.001
141:45:29.2* 0.010 0.003 0.018
*Systems aligned independently. Actual time of
abort guidance system alignment was 141:18:35.2


The performance of the abort sensor assembly of the abort guidance system was not as good as on previous missions but was within allowable limits. The accelerometers exhibited stable performance, but the Z-axis gyro drift rate change of 1.2 degrees per hour from the prelaunch value was about 30 percent greater than the expected shift. The expected and the actual shifts between preflight values and the first inflight calibration, and shifts between subsequent inflight calibrations are shown in table 8-IV.

TABLE 8-IV - ABORT GUIDANCE SYSTEM CALIBRATION COMPARISONS

Calibrations Three-sigma
capability
estimate
Actual gyro drift rate, deg/hr
X axis Y axis Z axis
First inflight minus
pre-installation
±0.91 0.08 -0.07 -1.2
Second inflight minus
first inflight
±0.63 -0.01 0.23 0.26
First surface minus
second inflight
±0.56 -0.02 -0.08 -0.43
Second surface minus
first surface
±0.55 0.00 -0.08 -0.21


Table 8-V is a sequence of events prior to and during the powered descent to the lunar surface. A command to abort using the descent engine was detected at a computer input channel at 104:16:07 (but was not observed at other telemetry points) although the crew had not depressed the abort switch on the panel. The crew executed a procedure using the engine stop switch and the abort switch which isolated the failure to the abort switch. Subsequently, the command reappeared three more times; each time, the command was removed by tapping on the panel near the abort switch. (For a discussion of the probable cause of this failure, see section 14.2.2.)


TABLE 8-V - SEQUENCE OF EVENTS DURING POWERED DESCENT

Elapsed time
from lift-off,
hr:min:sec
Time from
ignition,
min:sec
Event
107:51:18.66 -11:07.86 Landing radar on
107:52:46.66 -9:39.86 False data good indications
from landing radar
107:57:34.66 -4:51.86 Landing radar switched to low scale
107:58:13.80 -4:12.72 Start loading abort bit
work-around routine
108:02:19.12 -0:07.40 Ullage on
108:02:26.52 0:00.00 Ignition
108:02:53.80 +0:27.28 Manual throttle-up
to full throttle position
108:04:49.80 +2:23.28 Manual target update (N69)
108:08:47.68 +6:21.16 Throttle down
108:08:50.66 +6:24.14 Landing radar to high scale
(circuit breaker cycle)
108:09:10.66 +6:44.14 Landing radar velocity data good
108:09:12.66 +6:46.14 Landing radar range data good
108:09:35.80 +7:09.28 Enable altitude updates
108:11:09.80 +8:43.28 Select approach phase program (P64)
108:11:10.42 +8:43.90 Start pitch over
108:11:51.60 +9:25.08 Landing radar redesignation enable
108:11:52.66 +9:26.14 Landing radar antenna to position 2
108:13:07.86 +10:41.34 Select attitude hold mode
108:13:09.80 +10:43.28 Select landing phase program (P66)
108:15:09.30 +12:42.78 Left pad touchdown
108:15:11.13 +12:44.61 Engine shutdown
(decreasing thrust chamber pressure)
108:15:11.40 +12:44.88 Right, forward, and aft pad touchdown


If the abort command is present after starting the powered descent programs, the computer automatically switches to the abort programs and the lunar module is guided to an abort orbit. To avoid the possibility of an unwanted abort, a work-around procedure was developed by ground personnel and was relayed to the crew for manual entry into the lunar module computer. Part one of the four-part procedure was entered into the computer just after the final attitude maneuver for powered descent. The remainder was accomplished after the increase to the full-throttle position. Part one consisted of loading the abort stage program number into the mode register in the erasable memory which is used to monitor the program number displayed to the crew. This did not cause the active program to change, but it did inhibit the computer from checking the abort command status bit. At the same time, it inhibited the automatic command to full-throttle position, automatic guidance steering, and it affected the processing of the landing radar data. Therefore, in order to reestablish the desired configuration for descent, the increase to full-throttle position was accomplished manually and then the second, third, and fourth parts of the procedure were entered into the computer. In order, they accomplished:
a. Setting a status bit to inform the descent program that throttleup had occurred and to re-enable guidance steering

b. Resetting a status bit which disabled the abort programs

c. Replacing the active program number back into the mode register so that landing radar data would be processed properly after landing radar lock-on

The abort capability of the primary guidance system was lost by use of this procedure. Therefore, it would have been necessary to use the abort guidance system if an abort situation had arisen.

Prior to powered descent maneuver ignition, the landing radar scale factor switched to low, which prevented acquisition of data through the first 400 seconds of descent. (For further discussion, refer to section 14.2.4.) The crew cycled the radar circuit breaker, which reset scaling to the high scale, and landing radar lock-on occurred at 22 486 feet. Figure 14-22 is a plot of slant range as measured by landing radar and as computed from primary guidance system state vectors. Figure 8-3 is a plot of altitudes computed by the abort and primary guidance systems and shows a 3400-foot update to the abort guidance system at the 12 000 foot altitude.

Figure 8-3 - Comparison of altitudes computed by abort and primary guidance systems during descent.


Throttle oscillations that had been noted on previous flights were not detected during the descent although some oscillation in the automatic throttle command was detected after descent engine manual shutdown. The reaction control system propellant consumption during the braking phase and approach phase programs was approximately half that seen on previous missions. Further discussion of these two areas will be provided in a supplement to this report.

While on the lunar surface, a test was performed to compute gravity using primary guidance system accelerometer data. The value of gravity was determined to be 162.65 cm/sec/sec.

Performance during the ascent from the lunar surface was nominal. The primary and abort systems and the powered flight processor data compared well throughout ascent. The ascent program in the onboard computer does not include targeting for a specific cutoff position vector; therefore, a vernier adjustment maneuver of 10.3 ft/sec was performed to satisfy the phasing conditions for a direct rendezvous with the command and service module.

Performance throughout rendezvous, docking, and the deorbit maneuver was also nominal. The velocity change imparted to the lunar module at jettison was minus 1.94, minus 0.05, and minus 0.10 ft/sec in the X, Y, and Z axes, respectively.

The abort guidance system functioned properly until the braking phase of the rendezvous with the command and service module when a failure caused the system to be down-moded to the standby mode and resulted in the loss of this system for the remainder of the mission. Another anomaly reported was a crack in the glass window of the address register on the data entry and display assembly. These anomalies are discussed in sections 14.2.5 and 14.2.6, respectively.

8.7 DESCENT PROPULSION

The descent propulsion system operation was satisfactory. The engine transients and throttle response were normal.

8.7.1 Inflight Performance

The duration of the powered descent firing was 764.6 seconds. A manual throttle-up to the full throttle position was accomplished approximately 26 seconds after the engine-on command. The throttle-down to 57 percent occurred 381 seconds after ignition, about 14 seconds earlier than predicted but within expected tolerances. Three seconds of the 14 are attributed to the landing site offset to correct for the downrange error in actual trajectory, and the remaining 11 seconds to a thrust increase of approximately 80 pounds at the full-throttle position.

8.7.2 System Pressurization

During the period from lift-off to 104 hours, the oxidizer tank ullage pressure decayed from 111 to 66 psia and the fuel tank ullage pressure decreased from 138 to 111 psia. These decays resulted from helium absorption into the propellants and were within the expected range.

The supercritical helium system performed as anticipated. The system pressure rise rates were 8.0 psi/hour on the ground and 6.2 psi/hour during translunar coast, which compare favorably with the preflight predicted values of 8.1 psi/hour and 6.6 psi/hr, respectively. During powered descent, the supercritical helium system pressure profile was well within the nominal 3-sigma pressure band, even though the pressure at ignition was about 50 Psi lower than anticipated.

8.7.3 Gaging System Performance

The gaging system performance was satisfactory throughout the mission. The low-level quantity light came on approximately 711 seconds after ignition, and was most probably triggered by the point sensor in oxidizer tank 2. Engine cutoff occurred 53 seconds after the low-level signal, indicating a remaining firing-timeto-depletion of 68 seconds. Using probe data to calculate remaining firing time gave approximately 70 seconds remaining. This is within the accuracy associated with the propellant quantity gaging system.

The new propellant slosh baffles installed on Apollo 14 appear to be effective. The propellant slosh levels present on Apollo 11 and 12 were not observed in the special high-samplerate gaging system data of this mission.

8.8 ASCENT PROPULSION

The ascent propulsion system duty cycle consisted of two firings - the lunar ascent and the terminal phase initiation. Performance of the system for both firings was satisfactory. Table 8-VI is a summary of actual and predicted performance during the ascent maneuver. The duration of engine firing for lunar ascent was approximately 432 seconds, and for terminal phase initiation, 3 to 4 seconds. A more precise estimate of the terminal phase initiation firing time is not available because the firing occurred behind the moon and no telemetry data were received. System pressures were as expected both before and after the terminal phase initiation maneuver and crew reports indicate that the maneuver was nominal.

TABLE 8-VI - STEADY-STATE PERFORMANCE DURING ASCENT

Parameter 10 seconds after ignition 400 seconds after ignition
Predicted* Measured** Predicted* Measured**
Regulator outlet pressure, psia 184 182 184 181
Oxidizer bulk temperature, F 70.0 69.4 69.0 69.4
Fuel bulk temperature, F 70.0 69.8 69.8 69.4
Oxidizer interface pressure, psia 170.5 168 169.7 167
Fuel interface pressure, psia 170.4 169 169.7 167
Engine chamber pressure, psia 123.4 121 123.2 120
Mixture ratio 1.607 - 1.598 -
Thrust, lb 3502. - 3468. -
Specific impulse, sec 310.3 - 309.9 -
* Preflight prediction based on acceptance test data and assuming nominal system performance.
** Actual flight data vith no adjustments.


No oscillations were noted during flight in either helium regulator outlet pressure measurement. Oscillations in the outlet pressure of 6 to 19.psi have been noted in previous flight data. Also, oscillations of a similar nature and approximately twice that magnitude were noted during preflight checkout of the ascent propulsion system class I secondary helium regulator. However, during flight, control is maintained, normally, by the class I primary regulator.

8.9 ENVIRONMENTAL CONTROL AND CREW STATION

Performance of the environmental control system was satisfactory throughout the mission. Glycol pump noise, a nuisance experienced on previous missions, was reduced below the annoyance level by a muffler on the pump system. Although the water separator speed was higher than expected much of the time, the separator removed water adequately and there were no problems with water condensation or cabin humidity.

Because of water in the suit loop on Apollo 12 (ref. 1), a flow restrictor had been instal-led in the primary -lithium hydroxide cartridges to reduce the gas flow in the suit loop and, thereby, reduce water separator speed below 3600 rpm. (Separator speed is a function of the water mass to be separated and the gas flow.) However, the water separator speed was above 3600 rpm 'while the suit was operated in the cabin mode (helmets and gloves removed). The high speed when in the cabin mode resulted from low moisture inputs from the crew (approximately 0.14 lb/hr) and a high gas flow caused by low back pressure which, in turn, developed from a low pressure drop across the suit.

During preparations for the first extravehicular activity, the transfer hose on the urine collection transfer assembly was kinked. The kink was eliminated by moving the hose to a different position.

The crew repeatedly had trouble getting the lunar module forward window shades to remain in their retainers. The shades had been processed to reduce the curl and prevent cracking, a problem experienced on previous flights. In reducing the curl, the diameter of the rolled shades was increased so that the shades would not fit securely in the retainers. For Apollo 15, the shades will be fabricated to permit them to be rolled small enough to be held securely by the retainers.

The interim stowage assembly could not be secured at all times because the straps could not be drawn tight enough to hold. This problem resulted from stretch in the fabric and in the sewing tolerances. In the future, more emphasis will be placed upon manufacturing fit checks and crew compartment fit checks to assure that the problem does not recur.

8.10 EXTRAVEHICULAR MOBILITY UNIT

Performance of the extravehicular mobility unit was very good during the entire lunar stay. Oxygen, feedwater, and power consumption (section 8.11-7) allowed each extravehicular period to be extended approximatelY 30 minutes with no depletion of contingency reserves. Comfortable temperatures were maintained using the diverter valve in the minimum position throughout most of both extravehicular activities.

Preparations for the first extravehicular activity proceeded on schedule with few exceptions. The delay in starting the first extravehicular activity occurred while the portable life support system power was on, resulting in battery power being the limiting consumable in determining the extravehicular stay time.

Oxygen consumption of the Lunar Module Pilot during the first extravehicular activity was one-third higher than that of the Commander. Telemetry data during the Lunar Module Pilot's suit integrity check indicated a pressure decay rate of approximately 0.27 psi/min; a rate of 0.30 psi/ min is allowable. In preparation for the second extravehicular activity, special attention was given to cleaning and relubricating the Lunar Module Pilot's pressure garment assembly neck and wrist ring seals in an effort to lower the extravehicular mobility unit leak rate. A 0.22 psi/min pressure decay rate was reported by the Lunar Module Pilot prior to the second extravehicular activity. Postflight unmanned leak rate tests on the Lunar Module Pilot's pressure garment assembly show no significant increase in leakage.

Just prior to lunar module cabin depressurization for the second extravehicular activity, the Lunar Module Pilot reported a continuous force in his right extravehicular glove wrist pulling to the left and down. A more detailed discussion is given in section 14.3.2. The extravehicular activity started and was completed without any reported difficulty with the glove.

8. 11 CONSUMABLES

On the Apollo 14 mission, all lunar module consumables remained well within red line limits and were close to predicted values.

8.11.1 Descent Propulsion System

Propellant.- The quantities of descent propulsion system propellant loading in the following table were calculated from readings and measured densities prior to lift-off.

TABLE - DESCENT PROPULSION SYSTEM PROPELLANTS

Condition Actual quantity, lb
Fuel Oxidizer Total
Loaded 7072.8 11 344.4 18 417.2
Consumed 6812.8 10 810.4 17 623.2
Remaining at engine cutoff

Total
Usable


260.0
228.0


534.0
400.0


794.0
628.0



Supercritical helium.- The quantities of supercritical helium were determined by computation utilizing pressure measurements and the known volume of the tank.

TABLE - DESCENT PROPULSION SYSTEM SUPER-CRITICAL HELIUM

Condition Quantity, lb
Actual Predicted
Loaded 48.5 -
Consumed

42.8

39.2
(40 .8)*
Remaining at touchdown

5.7

9.3
(7.7)*
* Adjusted prediction
to account for longer-than-planned firing duration



8.11.2 Ascent Propulsion System

Propellant, Ascent propulsion system total propellant usage was within approximately 1 percent of the predicted value. The loadings in the following table were determined from measured densities prior to launch and from weights of off-loaded propellants.

TABLE - ASCENT PROPULSION SYSTEM PROPELLANTS

Condition Actual quantity, lb Predicted
quantity , lb
Fuel Oxidizer Total
Loaded 2007.0 3218.2 5225.2  
Total consumed 1879.0 3014.0 4893.0 4956.0
Remaining at lunar
module jettison
128.0 204.2 332.2 265.8



Helium.- The quantities of ascent propulsion system helium were determined by pressure measurements and the known volume of the tank.

TABLE - ASCENT PROPULSION SYSTEM HELIUM

Condition Actual
quantity, lb
Loaded 13.4
Consumed 8.8
Remaining at
lunar module impact
4.6



8.11.3 Reaction Control System Propellant

The reaction control system propellant consumption was calculated from telemetered helium tank pressure histories using the relationships between pressure, volume, and temperature.

TABLE - REACTION CONTROL SYSTEM PROPELLANTS

Condition Actual, lb Predicted, lb
Fuel Oxidizer Total
Loaded

System A
System B

Total


108
108

216


209
209

418





634





633
Consumed to

Docking
Impact
   

260
378


283
393
Remaining at lunar impact     256 240



8.11.4 Oxygen

The oxygen tank was not loaded to the nominal 2730 psia used for previous missions because of a possible hydrogen embrittlement problem with the descent stage oxygen tank. Launch pressure for the tank was an indicated 2361 psia.

TABLE - LUNAR MODULE OXYGEN QUANTITIES

Condition Actual
quantity, lb
Predicted
quantity, lb
Loaded (at lift-off)
Descent stage
Ascent stage
Tank 1
Tank 2

Total

42.3

2.4
2.4

47.1
 
Consumed
Descent stage
Ascent stage
Tank 1
Tank 2

Total

24.9

(*)
0

-

23.9

1.1
0

25.0
Remaining in descent stage at
lunar lift-off
17.4 18.4
Remaining at docking
Tank 1
Tank 2

Total

(*)
2.4

-

1.3
2.4

3.7
* Consumables data are not available because the
tank 1 pressure transducer malfunctioned before launch.



8.11.5 Water

In the following table, the actual quantities loaded and consumed are based on telemetered data.

TABLE - LUNAR MODULE WATER QUANTITIES

Condition Actual
quantity, lb
 
Loaded (at lift-off)
Descent stage
Ascent stage
Tank 1
Tank 2

Total

255.5

42.5
42.5

340.5
 
Consumed
Descent stage (lunar lift-off)
Ascent stage (docking)
Tank 1
Tank 2

Total

200.9

6.0
5.8

212.7

190.9

6.2
6.2

203.3
Consumed
Descent stage (lunar lift-off)
Ascent stage (impact)
Tank 1
Tank 2

Total

200.9

14.4
14.9

230.2


-

-
-

-

Remaining in descent stage at
lunar lift-off
54.6 59.1
Remaining in ascent stage at impact
Tank 1
Tank 2

Total

28.1
27.6

55.7

-
-

-



8.11.6 Electrical Power

The total battery energy usage is given in the following table. Preflight predictions versus actual usage were within 3 percent.

TABLE - LUNAR MODULE BATTERY ENERGY

Batteries Available
power,
Amp-hrs
Electrical power consumed, Amp-hrs
Actual Predicted
Descent 1600 1191 1220
Ascent 592 128 125



8.11.7 Extravehicular Mobility Unit

Oxygen, feedwater and power consumption of the extravehicular mobility unit for both extravehicular periods are shown in the following table.

TABLE - EMU CONSUMABLES

Condition Commander Lunar Module Pilot
Actual Predicted Actual Predicted
First extravehicular activity
Time, min 288 255 288 255
Oxygen, lb
Loaded
Consumed
Remaining

1.31
0.70
0.61

1.31
0.97
0.34

1.31
1.02
0.29

1.31
0.97
0.34
Feedwater, lb
Loaded
Consumed
Remaining

8.59
4.85
3.74

8.55
7.08
1.47

8.66
5.71
2.95

8.55
7.08
1.47
Power, W-h
Initial charge
Consumed
Remaining

282
228
54

282
223
59

282
237
45

282
223
59
Second extravehicular activity
Time, min 275 255 275 255
Oxygen, lb
Loaded
Consumed
Remaining

1.26
0.86
0.40

1.31
1.02
0.29

1.26
0.96
0.30

1.31
1.02
0.29
Feedwater, lb
Loaded
Consumed
Remaining

8.80
6.43*
2.37*
8.55
7.55
1.0
8.80
7.13*
1.67*
8.55
7.55
1.0
Power, W-h
Initial charge
Consumed
Remaining

282
225
57

282
225
57

282
222
60

282
225
57
*Estimate based on extravehicular mobility unit source heat predictions because
portable life support system feedwater weight was
not taken following the second extravehicular activity.

Chapter 9 - Pilot's Report Table of Contents Apollo 14 Journal Index