7.0 COMMAND AND SERVICE MODULE PERFORMANCE

7.1 STRUCTURAL AND MECHANICAL SYSTEMS

Structural loads on the spacecraft during all phases of the mission were within design -limits. The predicted and calculated loads at liftoff,, in the region of maximum dynamic pressure, at the end of first stage boost, and during staging were similar to those of previous missions. Command module accelerometer data prior to S-IC center engine cutoff indicate a sustained 5-hertz longitudinal oscillation with an amplitude of 0.179, which is similar to that measured during previous flights. Oscillations during the S-II boost phase had a maximum measured amplitude of less than 0.06g at a frequency of 9 hertz. The amplitudes of both oscillations were within acceptable structural design limits.

Six attempts were required to dock the command and service module with the lunar module following translunar injection. The measured rates and indicated reaction control system thruster activity during the five unsuccessful docking attempts show that capture should have occurred each time. The mechanism was actuated and inspected in the command module following docking. This investigation indicated that the probe mechanical components were functioning normally. Subsequent undocking and docking while in lunar orbit were normal. The probe was returned for postflight analysis. The docking anomaly is discussed in detail in section 14.1.1.

7.2 ELECTRICAL POWER

7.2.1 Power Distribution

The electrical power distribution system performed normally except for two discrepancies. Prior to entry, when the bus-tie motor switches were operated to put the entry batteries on the main busses, battery C was not placed on main bus B. This anomaly was discovered by the data review after the flight. Postflight continuity checks revealed that the circuit breaker tying battery C to main bus B was inoperative. This anomaly is described in section 14.1-7.

The second discrepancy occurred during entry. Procedures call for main bus deactivation, at 800 feet altitude, by opening the bus tie motor switches. The crew reported that operation of the proper switches did not remove power from the buses. The buses were manually deactivated, after landing, by opening the in-line circuit breakers on Panel 275 (a normal procedure). Review of data indicated and postflight tests confirmed that the motor switch which tied battery A to main bus A was inoperative. This anomaly is described in section 14.1.6.

7.2.2 Fuel Cells

The fuel cells were activated 48 hours prior to launch, conditioned for 4 hours, and configured with fuel cell 2 on the line supplying a 20-ampere load as required in the countdown procedure. Fuel cells 1 and 3 remained on open circuit until 5 hours prior to launch. At launch, fuel cell 1 was on main bus A with fuel cell 2, and fuel cell 3 was on main bus B. This configuration was maintained throughout the flight. Initially, the load variance was approximately 5 amperes, but it stabilized to 3 or 4 amperes early in the flight. This is normal and typical of other flights.

All fuel cell parameters remained within normal operating limits and agreed with predicted flight values. As expected, the fuel cell 1 condenser-exit temperature exhibited a periodic fluctuation about every 6 minutes throughout the flight. This zero-gravity phenomenon was similar to that observed on all other flights and has no effect on fuel cell performance (ref. 6).

The fuel cells supplied 435 kW-h of energy at an average current of 23 amperes per fuel cell and a mean bus voltage of 29 volts during the mission.

7.2.3 Batteries

The command and service module entry and pyrotechnic batteries performed normally. Entry batteries A and B were both charged once at the launch site and five times during flight with nominal charging performance. Load sharing and voltage delivery were satisfactory during each of the service propulsion firings. The batteries were essentially fully charged at entry.

7.3 CRYOGENIC STORAGE

Cryogenics were satisfactorily supplied to the fuel cells and to the environmental control system throughout the mission. The configuration changes made as a result of the Apollo 13 oxygen tank failure are described in appendix A. A supplemental report giving details of system performance will be issued at a later date (appendix E).

During preflight checkout of the oxygen system, the singleseat check valve for tank 2 was found to have failed in the open position and was replaced with an in-line double-seat valve. During flight, this valve allowed gas leakage into tank 2 from tank 3. The purpose of this valve is primarily to isolate tank 2 from the remainder of the system should tank 2 fail. Thus, it was qualified at a reverse differential. pressure of 60 psid. This is significantly higher than that normally experienced during a flight. Tests have been conducted to characterize the nature of the check valve leakage at low pressure differential and show that this situation is not detrimental to operation under abnormal as well as normal conditions.

Two flow tests on the oxygen system were conducted during flight. One was to demonstrate the capability of the system to support additional flow requirements for extravehicular activities. The other was to determine the heater temperature while operating with the oxygen density less than 20 percent. The intent of these two tests was met and favorable results were obtained although test procedures were modified because of time constraints. The oxygen system is capable of supporting the anticipated requirements for Apollo 15 and subsequent missions. The lowdensity flow test indicated that the oxygen system can provide required flow rates at low densities and the data obtained provides for a more accurate assessment of heater operating temperature.

Consumable quantities in the cryogenic storage system are discussed in section 7.10.3.

7.4 COMMUNICATIONS EQUIPMENT

The communications system satisfactorily supported the mission except for the following described conditions.

The high-gain antenna failed to acquire and track properly at various times during the mission. The problems occurred during the acquisition of signal rather than after acquisition. In this regard, the problem is different from those experienced during Apollo 12 and 13 where the highgain antenna lost lock or failed to track after acquisition. This is discussed in further detail in section 14.1.2.

From just prior to lunar lift-off through terminal phase initiation, the VHF system performance was marginal. Voice communications were weak and noisy, and the VHF ranging performance was erratic and erroneous. The voice communications problem is not related to the VHF problems experienced on previous missions where they were determined to be procedural errors. Switching antennas in the command and service module and elimination of the ranging signal did not clear up the problems. The problems are believed to have been caused by equipment malfunction, but the source has not been isolated to a particular component of the total system. Section 14.1.4 contains a detailed discussion of this anomaly.

7.5 INSTRUMENTATION

The instrumentation system functioned normally throughout the mission except for the loss of the reaction control system quad B oxidizer manifold pressure measurement during separation of the command and service module from the launch vehicle. The most probable cause of the failure was a break of the signal or power leads initiated by the pyrotechnic shock associated with the spacecraft/launch vehicle adapter panel separation. Since this is the only failure of four measurements of this type on each of eight flights, the pyrotechnic shock is not considered a problem for normal elements of the instrumentation circuit. Further, redundant measurements are available to permit determination of the required data. Consequently, no corrective action is required.

7.6 GUIDANCE, NAVIGATION, AND CONTROL

Attitude control was nominal throughout the mission including all periods of passive thermal control, cislunar navigation, as well as photography and landmark tracking from lunar orbit. The stability of the inertial measurement unit error parameters was excellent. The only anomaly in the guidance, navigation and control system was failure of the entry monitor system 0.059 light to illuminate. This is discussed in section 14.1-5.

Because of inclement weather, the lift-off was delayed for the first time in the Apollo program. This required the flight azimuth to be changed from 72 degrees to 75-56 degrees and the platform to be realigned accordingly. A comparison of command and service module and S-IVB navigation data indicated satisfactory performance during the launch phase. Insertion errors were plus 7.02, plus 61.02, and minus 7.50 ft/sec in the X, Y, and Z axes, respectively. These errors were comparable to those observed on other Apollo launches. The only significant error was in the Y-a.xis velocity caused by a prelaunch azimuth alignment error of 0.14 degree due to one-sigma gyrocompassing inaccuracies. Table 7-I is a summary of preflight inertial measurement unit error parameters after its installation in the command module. An update to the inertial parameters was performed at approximately 29 hours. The three accelerometer biases were updated to minus 0.32, plus 0.12 and minus 0.13 cm/sec2' and the X-gyro null bias drift was updated to plus 0.4 meru (milli earth-rate units).

TABLE 7-I - INERTIAL COMPONENT PREFLIGHT HISTORY - COMMAND MODULE

Error Sample
mean
Standard
deviation
No. of
samples
Countdown
value
Flight
load
Inflight
performance
Accelerometers
X - Scale factor error, ppm

Bias , cm/sec/sec
-444

-0.23
58

0.13
8

8
-500

-0.31
-370

-0.23
-

-0.34
Y - Scale factor error, ppm

Bias , cm/sec/sec
-441

0.05
49

0.07
8

8
-505

0.13
-500

0.04
-

0.09
Z - Scale factor error, ppm

Bias , cm/sec/sec
-278

-0.29
49

0.07
8

8
-320

-0.18
-310

-0.29
-

-0.18
Gyroscopes
X - Null bias drift, meru

Acceleration drift,
spin reference axis,
meru/g

Acceleration drift,
input axis,
meru/g
0.9



3.0



1.7
0.6



2.0



1.5
8



8



8
1.8



4.9



-1.6
2.5



1.0



0.0
0.0*



-



-
Y - Null bias drift, meru

Acceleration drift,
spin reference axis,
meru/g

Acceleration drift,
input axis,
meru/g
-3.4



3.2



-9.9
0.8



1.5



4.5
8



8



16
-4.2



3.8



-9.7
-3.4



3.0



-5.0
1.7



-



-
Z - Null bias drift, meru

Acceleration drift,
spin reference axis,
meru/g

Acceleration drift,
input axis,
meru/g
1.6



-3.1



43.8
0.9



1.0



6.4
8



8



8
2.5



-2.4



54.1
1.6



-3.9



40.0
0.0



-



-
*lnflight performance average before update was minus 2.0.



The first platform realignment was performed after insertion and agreed with the predicted alignment errors due to prelaunch azimuth errors. Table 7-II is a summary of significant parameters during each of the platform realignments.



TABLE 7-II - COMMAND AND SERVICE MODULE PLATFORM ALIGNMENT SUMMARY

Time,
hr:min
Program
option*
Star used Gyro torquing angle, deg Star angle
difference, deg
Gyro drift, meru Comments
X Y Z X Y Z
000:58 REFSMMAT 22 Regulus, 24 Gienah 0.085 0.010 0.166 0.00       Launch orientation
006:40 REFSMMAT 17 Regor, 14 Canopus 0.127 -0.060 -0.011 0.00 -1.4 +0.7 -0.1 Launch orientation
014:13 REFSMMAT 31 Arcturus, 35 Hasalhague 0.271 -0.127 -0.036 0.01 -2.5 1.2 -0.3 Passive thermal control
029:20 REFSMMAT 20 Dnoces, 23 Denebola 0.449 -0.130 0.082 0.01 -2.0 0.6 0.4 Passive thermal control
040:11 REFSMMAT 1 Alpheratz, 40 Altair -0.039 -0.221 0.046 0.00 0.2 1.4 0.3 Passive thermal control
053:11 REFSMMAT 20 Dnoces, 23 Denebola 0.006 -0.129 0.052 0.00 -0.0 0.7 0.3 Passive thermal control
059:41 REFSMMAT 13 Capella, 3 Navi -0.073 -0.093 0.033 0.00 0.8 1.1 0.4 Passive thermal control
076:52 REFSMMAT 23 Denebola, 32 Alpheeca 0.056 -0.262 0.038 0.00 -0.2 1.0 0.1 Passive thermal control
079:39 REFSMMAT 27 Alkaid, 35 Rasalhague -0.007 -0.045 0.010 0.00 0.2 1.1 0.2 Passive thermal control
084:09 REFSMMAT 30 Menkent, 35 Rasalhague 0.001 -0.055 0.002 0.01 -0.2 1.2 -0.5 Landing site orientation
086:10 REFSMMAT 16 Procyon, 17 Regor -0.050 -0.070 -0.045 0.01 1.7 2.3 -1.5 Landing site orientation
088:05 REFSMMAT 16 Procyon, 20 Dnoces -0.031 0.002 0.027 0.01 1.1 0.1 0.9 Landing site orientation
101:24 REFSMMAT 17 Regor, 30 Menkent 0.073 -0.229 0.000 0.00 -0.4 1.1 0.0 Landing site orientation
105:09 REFSMMAT 40 Altair, 42 Peacock 0.030 -0.038 0.028 0.01 -0.6 0.7 0.2 Landing site orientation
109:12 REFSMMAT 34 Atria, 37 Nunki -0.012 -0.043 0.003 0.01 0.2 0.7 0.0 Landing site orientation
117:08 REFSMMAT 22 Regulus, 27 Alkaid 0.021 -0.105 0.055 0.02 -0.2 0.9 0.5 Landing site orientation
119:27 REFSMMAT 12 Rigel, 21 Alphard -0.027 -0.065 0.018 0.00 1.3 1.9 0.5 Launch orientation
131:19 REFSMMAT 10 Mirfak, 12 Rigel -0.036 -0.157 0.091 0.01 0.3 1.2 0.7 Launch orientation
137:18 REFSMMAT 6 Acamar, 14 Canopus -0.002 -0.166 -0.005 0.00 0.0 1.8 -0.1 Launch orientation
140:53 REFSMMAT 31 Arcturus, 30 Menkent 0.079 -0.006 -0.001 0.00 -1.3 0.1 -0.0 Launch orientation
146:58 REFSMMAT 24 Gienah, 31 Arcturus 0.018 -0.091 0.050 0.00 -0.2 1.0 0.5 Launch orientation
150:17 REFSMMAT 4 Achernar, 34 Atria 0.037 -0.106 -0.043 0.01 -0.7 2.1 0.9 Transearth injection orientation
163:49 REFSMMAT 11 Aldebaran, 16 Procyon 0.046 -0.174 0.017 0.00 -0.2 0.8 0.1 Passive thermal control orientation
186:34 REFSMMAT 25 Acrux, 42 Peacock 0.040 -0.460 0.076 0.00 -0.1 1.3 0.1 Passive thermal control orientation
192:14 REFSMMAT 41 Dabih, 34 Atria -0.038 -0.104 -0.003 0.01 0.1 1.2 0.0 Passive thermal control orientation
196:58 REFSMMAT 17 Regor, 40 Altair -0.009 -0.109 0.038 0.01 0.1 1.5 0.5 Passive thermal control orientation
208:11 REFSMMAT 25 Acrux, 33 Antares 0.071 -0.161 0.026 0.01 -0.4 1.0 0.2 Passive thermal control orientation
212:59 REFSMMAT 16 Procyon, 23 Denebola -0.049 -0.010 0.014 0.01 0.7 0.1 0.2 Passive thermal control orientation
213:11 Preferred 23 Denebola, 16 Procyon 0.021 0.002 -0.036 0.01 -1.0 -1.0 -1.6 Entry orientation
214:39 REFSMMAT 30 Menkent, 37 Nunki 0.039 -0.040 -0.069 0.00 -1.8 1.8 -3.2 Entry orientation
* 1 - Preferred; 2 - Nominal; 3 - REFSMMAT; 4 - Landing site                



Spacecraft dynamics during separation from the S-IVB were very small. Spacecraft dynamics during each docking attempt were small and comparable to those seen on previous Apollo missions. Figure 7-1 is a time history of significant control system parameters during each docking attempt.

Figure 7-1.-History of control system parameters during multiple docking attempts.


Performance during each of the seven service propulsion system maneuvers was nominal. Trimming of residual velocity errors was performed only after the circularization and transearth injection maneuvers. Table 7-III is a summary of significant control system parameters for each of the maneuvers. The second midcourse correction was accomplished with a minimumimpulse service propulsion system maneuver in order to conserve reaction control propellant. This was the first service propulsion system minimum-impulse maneuver perf6rmed during a lunar mission. The third midcourse correction was performed with the reaction control system.

TABLE 7-III - GUIDANCE AND CONTROL MANEUVER SUMMARY

Parameter Maneuver
First
midcourse
correction
Second
midcourse
correction
Lunar orbit
insertion
Descent orbit
insertion
Lunar orbit
circularization
First
plane change
Transearth
injection
midcourse
correction
(performed with RCS)
Time

 Ignition, hr:min:sec
 Cutoff, hr:min:sec
 Duration, min:sec


30:36:07.91
30:36:18.10
0:10.19


76:58:11.98
76:58:12.63
0:00.65


81:56:40.70
82:02:51.54
6:l0.84


86:10:52.97
86:11:13.78
0:20.81


105:11:46.11
105:11:50.13
0:04.02


117:29:33.17
117:29:51.67
0:18.50


148:36:02.30
148:38:31.53
2:29.23


165:34:56.69

0:03.00
Velocity gained, ft/sec,
inertial coordinates
(desired/actual)

 X
 Y
 Z




+11.0/+10.9
+63.1/+63.3
+30.9/+30.9




-1.8/-1.9
+0.3/+0.2
+3.3/+3.4




+1957.9/+1958.2
-2301.0/-2301.2
+80.0/+79.9




+185.3/+185.7
-51.4/-52.5
-73.0/-73.2




-76.8/-74.9
-11.1/-l0.6
-9.6/-9.3




-74.5 /-74.4
+188.1 /+188.0
-310.1/-310.9




-3284.7/-3285.4
+236.3/+236.6
-1061.3/-1061.8




-0.5/-0.7
+0.2/0.0
+0.1/0.0
Velocity residual, ft/sec
body coordinates
(+ indicates underburn)

 X
 Y
 Z
 Entry monitor system




+0.3
0.0
-0.1
+0.3




+0.3
0.0
0.0
+0.5




+0.3
0.0
0.0
-0.3




+0.6
+0.2
0.0
-0.2




-1.0
0.0
+0.5
+0.4




+0.6
+0.4
+0.2
+1.2



+0.1
+0.8
-0.3
+2.5




+0.2
+0.2
+0.1
0.0
Engine gimbal position,
 deg

 Initial
   Pitch
   Yaw

 Maximum excursion
   Pitch
   Yaw

 Steady-state
   Pitch
   Yaw

 Cutoff
   Pitch
   Yaw




+1.00
-0.18


+0.32
-0.47


+1.00
-0.18


+1.00
-0.26




+0.87
-0.24


+0.05
0.09


N/A
N/A


+0.92
-0.35




+0.87
-0.26


+0.49
+0.49


+1.14
-0.26


+1.63
-0.65




+1.50
-0.60


+0.27
-0.30


+1.59
-0.60


+1.72
-0.60




-0.75
+0.24


-1.92
+1.61


-0.71
+0.12


-0.71
+0.07




-0.88
+0.20


-2.14
+1.53


-0.68
+0.16


-0.62
-0.05




-0.66
+0.12


-2.10
+1.27


-0.53
-0.26


-0.62
-1.62
N/A
Max. rate excursion,
deg/sec

 Pitch
 Yaw
 Roll



-0.12
-0.12
±0.08



0.0
0.0
0.0



-0.16
+0.16
+0.20



+0.28
+0.20
+0.12



+1.23
-0.68
-0.59



+1.42
-1.12
-0.72



+1.32
-1.32
-1.86
N/A
Max. attitude error, deg

 Pitch
 Yaw
 Roll


-0.15
-0.22
-1.31


0.0
-0.04
0.0


+0.16
-0.15
+5.00


-0.16
-0.08
-0.60


-0.31
-0.14
-0.84


+0.25
-0 .26
-3.78


+0.24
-0.31
±5.00
N/A



During the translunar phase, a series of star-horizon measurements were taken to establish the precise location of the earth horizon. This was done in preparation for a cislunar navigation exercise to be performed during the transearth phase.

The command and service module combination was separated from the lunar module after the descent orbit insertion maneuver. Command and service module circularization and planechange maneuvers were then performed, and the Command Module Pilot accomplished a series of photographic and landmark tracking operations. For the first time, rate-aided optics were available to assist the crew in making optical sightings.

The sextant and VHF ranging data were used to track the lunar module after the vernier adjustment maneuver following ascent from the lunar surface. Unacceptable VHF ranging data were received in the interval between lunar module insertion and the terminal phase initiation maneuver; however, the data received during the final phase of rendezvous were good. For a detailed discussion of rendezvous, see section 6.2-3. For a discussion of the VHF ranging anomaly, see section 14.1.4.

Only one midcourse correction was required on the return trip to meet the entry interface conditions. Cislunar navigation was performed during the transearth phase to simulate returning to earth with no communications. Accuracy of the onboard navigation techniques was demonstrated but the crew commented that the computer/crew operational interface could be improved by incorporating a recycle feature in the cislunar navigational sighting program.

The command module was separated from the service module at 215:32:42 and the normal pitch-down disturbance was observed. The entry monitor system 0.05g light did not illuminate within the allowed 3 seconds after the predicted time for 0.05g. The crew started the system manually according to the checklist. Refer to section 14-1.5 for further discussion of this anomaly. Table 7-IV is a summary of entry monitor system nullbias tests performed during the mission. Accelerometer stability and performance were excellent.

The primary guidance system guided the command module to a landing at 27 degrees 0 minutes 45 seconds south latitude and 172 degrees 39 minutes 30 seconds west longitude, which is 0.62 mile from the targeted landing point.

TABLE 7-IV - RESULTS OF ENTRY MONITOR SYSTEM NULL BIAS TESTS

Test
(each of 100 s duration)
1 2 3 4 5 6 7 8 9
Time 01:50:00 09:34:50 29:11:20 58:28:00 75:59:00 79:45:00 84:31:00 118:20:00 165:15:00
Entry monitor system reading
 at start of test, ft/sec
-100 -100 -100 -100 -100 -100 -100 -100 -100
Entry monitor system reading
 at end of test, ft/sec
-99.5 -99.4 -99.6 -98.9 -98.4 -98.5 -99.4 -98.5 -99.0
Differential velocity bias, ft/sec
 (Count up is positive bias)
+0.5 +0.6 +0.4 +1.1 +1.6 +1.5 +0.6 +1.5 +1.0
Null bias, ft/sec/sec +0.005 +0.006 +0.004 +0.011 +0.016 +0.015 +0.006 +0.015 +0.010



7.7 REACTION CONTROL SYSTEMS

7.7.1 Service Module

Performance of the service module reaction control was normal throughout the mission. All telemetry parameters stayed within nominal limits throughout the mission with the exception of the quad B oxidizer manifold pressure. This measurement was lost when the command and service module separated from the SIVB. The quad B helium and fuel manifold pressures were used to verify proper system operation. Total propellant consumption for the mission was 102 pounds less than predicted; however, propellant consumption during transposition, docking and extraction was about 60 pounds more than planned because of the additional maneuvering associated with the docking difficulties. The propellant margin deficiency was recovered prior to lunar orbit insertion, and nominal margins existed during the remainder of the mission. Consumables information is contained in section 7.10.2.

7.7.2 Command Module

The command module reaction control systems performed satisfactorily. Both systems 1 and 2 were activated during the command module/service module separation sequence. Shortly after separation, system 2 was disabled and system 1 was used for the remainder of entry. All telemetry data indicated nominal system performance throughout the mission. Consunables information is contained in section 7.10.2.

7.8 SERVICE PROPULSION SYSTEM

Service propulsion system performance was satisfactory based on the steady-state performance during all firings. The steadystate pressure data, gaging system data, and velocities gained indicated essentially nominal performance. The engine transient performance during all starts and shutdowns was satisfactory. Nothing in the flight data or postflight analysis indicated combustion instability or the cause of the slight hum or buzzing noise reported by the pilot (ref. 9.13).

The propellant utilization and gaging system provided nearideal propellant utilization. The unbalance at the end of the transearth injection firing was reported by the crew to be 40 lbm, decrease, which agrees well with telemetry values.

During the Apollo 9, 10, 11, and 12 missions, the service propulsion system mixture ratio was less than expected, based on static firing data. The predicted flight mixture ratio for this mission was based on previous flight data to more closely simulate the expected mixture ratio. To achieve the predicted mixture ratio at the end of the mission, the majority of the mission would have to be flown with the propellant utilization valve in the increase position. Consequently, the propellant utilization valve was in the increase position at launch.

Figure 7-2 shows the variance in fuel and oxidizer remaining at any instant during the lunar orbit insertion and transearth injection firings, as computed from the telemetry data, and the propellant utilization valve movements made by the crew. The preflight expectedyalues and propellant utilization movements are also shown. The service propulsion system propellant usage for the mission is discussed in section 7.10-1.

Figure 7-2 - Oxidizer unbalance during lunar orbit insertion and transearth injection firings.



7.9 ENVIRONMENTAL CONTROL AND CREW STATION

The environmental control system performed satisfactorily and provided a comfortable environment for the crew and adequate thermal control of the spacecraft equipment. The crew station equipment also satisfactorily supported the flight.

The environmental control system was used in conjunction with the cryogenic oxygen system to demonstrate the capability of providing oxygen at high flow rates such as those that will be required during extravehicular operations on future missions. A modified hatch overboard dump nozzle with a calibrated orifice was used to obtain the desired flow rate. The emergency cabin pressure regulator maintained the cabin pressure at approximately 4.45 psia. The test, scheduled to last 2-1/2 hours, was terminated after 70 minutes when the 100-psi oxygen manifold pressure decayed to about 10 psi. This was caused by opening of the urine overboard dump valve which caused an oxygen demand in excess of that which the oxygen restrictors were capable of providing. However, sufficient data were obtained during the test to determine the high-flow capability of the cryogenic oxygen system. (Also see section 7.3.)

Inflight cabin pressure decay measurements were made for the first time during most of the crew sleep periods to determine more precisely the cabin leakage during flight. Preliminary estimates indicate that the flight leakage was approximately 0.03 lb/hr. This leak rate is within design limits.

Partial repressurization of the oxygen storage bottles was required three times in addition to the normal repressurizations during the mission. This problem is discussed in section 14.1.8.

The crew reported several instances of urine dump nozzle blockage. Apparently the dump nozzle was clogged with frozen urine particles. The blockage cleared in all instances when the spacecraft was oriented so that the nozzle was in the sun. This anomaly is discussed further in section 14.1.3.

Intermittent communications dropouts were experienced by the Commander at 29 hours. The problem was corrected when the Commander's constant wear garment electrical adapter was replaced. The anomaly is discussed further in section 14-3.4.

A vacuum cleaner assembly and cabin fan filter, used for the first time, along with the normal decontamination procedures eliminated practically all of the objectionable dust such as that present after the Apollo 12 lunar docking. The fans were operated for approximately 4 hours after lunar docking.

Sodium nitrate was added to the water buffer ampules to reduce system corrosion. This addition also allowed a reduction in the concentration of chlorine in the chlorine ampules. No objectionable taste was noted in the water. The crew reported some difficulty in inserting the buffer ampules into the injector. The ampules and injector are being tested to establish the cause of the problem. The crew also indicated that the food preparation unit leaked slightly after dispensing hot water. This problem is discussed further in section 14.1-7.

7.10 CONSUMABLES

The command and service module consumable usage during the Apollo 14 mission was well within the red line limits and, in all systems, differed no more than 5 percent from the predicted limits.

7.10.1 Service Propulsion Propellant

Service propulsion propellant loadings and consumption values are listed in the following table. The loadings were calculated from gaging system readings and measured densities prior to lift-off.

TABLE - SERVICE PROPULSION PROPELLANT LOADING AND COMSUMPTION

Condition Propellant, lb
Fuel Oxidizer Total
Loaded 15695.2 25061 40756.2
Consumed 14953.2 23900 38853.2
Remaining at command module/
service module separation
742 1161 1903
Usable at command module/
service module separation
596 866 1462



7.10.2 Reaction Control System Propellants

Service module - The propellant utilization and loading data for the service module reaction control system were as shown in the following table. Consumption was calculated from telemetered helium tank pressure histories and were based on pressure, volume, and temperature relationships.

TABLE - SERVICE MODULE REACTION CONTROL SYSTEM PROPELLANT LOADING AND CONSUMPTION

Condition Propellant, lb
Fuel Oxidizer Total
Loaded

  Quad A
  Quad B
  Quad C
  Quad D

  Total


110.1
109.9
110.4
109.7

440.1


225.3
225.2
226.5
223.5

900.5


335.4
335.1
336.9
333.2

1340.6
Usable loaded = amount loaded
minus amount trapped
with corrections made for gauging errors
1233
Consumed 250 476 726
Remaining at command module/
service module separation
507



Command module - The loading and utilization of command module reaction control system propellant was as follows. Consumption was calculated from pressure, volume and temperature relationships.



TABLE - COMMAND MODULE REACTION CONTROL SYSTEM PROPELLANT LOADING AND COMSUMPTION

Condition Propellant, lb
Fuel Oxidizer Total
Loaded

  System 1
  System 2

  Total


44.3
44.5

88.8


78.6
78.1

156.7


122.9
122.6

245.5
Usable loaded = amount loaded
minues the amount trapped and
with corrections made for gauging errors
210.0
Consumed

  System 1
  System 2

  Total


  41*
4

45
* Estimated quantity based on helium source pressure profile during entry.



7.10.3 Cryogenics

The total cryogenic hydrogen and oxygen quantities available at liftoff and consumed were as follows. Consunption values were based on quantity data transmitted by telemetry.

TABLE - TOTAL CRYOGENIC HYDROGEN AND OXYGEN LOADING AND CONSUMPTION

Condition Hydrogen, lb Oxygen, lb
Actual Planned Actual Planned
Available at lift-off

  Tank 1
  Tank 2
  Tank 3

  Total


26.97
26.55
-

53.52






53.52*


320.2
318.9
197.2

836.3






836.3*
Consumed

  Tank 1
  Tank 2
  Tank 3

  Total


19.12
19.14
-

38.26





38.62


119.3
113.8
163.4

396.5





412.1
Remaining at command module/
service module separation

  Tank 1
  Tank 2
  Tank 3

  Total



7.85
7.41
-

15.26



7.87
7.03
-

14.90



200.9
205.1
33.8

439.8



204.2
195.2
24.8

424.2
         
* Updated to lift-off values.



7.10.4 Water

The water quantities loaded, produced, and expelled during the mission are shown in the following table.

TABLE - COMMAND AND SERVICE MODULE WATER QUANTITIES

Condition Quantity, lb
Loaded (at lift-off)

  Potable water tank
  Waste water tank


28.5
32.4
Produced inflight

  Fuel cells
  Lithium hydroxide reaction
  Metabolic


342.3
21.0
21.0
Dumped overboard

  Waste tank dumping
  Urine and flushing


236.9
133.2
Evaporated up to command module/
service module separation

9.0
Remaining onboard at command module/
service module separation

  Potable water tank
  Waste water tank



29.7
36.4

Chapter 8 - Lunar Module Performance Table of Contents Apollo 14 Journal Index