6.0 COMMAND AND SERVICE MODULE PERFORMANCE.
6.1 STRUCTURAL AND MECHANICAL SYSTEMS
Command module accelerometer data indicated a sustained 5-hertz
longitudinal oscillation of 0.35g peak-to-peak amplitude prior to first stage
center engine cutoff. Similar oscillations have occurred on previous Apollo
flights and are within acceptable structural design limits . Oscillations
measured during second and third stage boost were less than 0.05g peak
amplitude in any direction and were not structurally significant.
Translunar docking loads were higher than those of previous missions
(see sec. 7-1).
Main parachute deployment for earth landing, beginning at
approximately 10 500 feet, was normal. However, at approximately 6000 feet,
one of the three main parachutes was observed to have collapsed. Details of
this anomaly are reported in section 14.1.9.
6.2 ELECTRICAL POWER AND FUEL CELLS
The electrical power system batteries and fuel cells performed
satisfactorily throughout the mission.
The entry, auxiliary, and pyrotechnic batteries performed normally.
Entry batteries A and B were charged nine times during flight (battery A
- 4 times; battery B - 5 times). Load sharing and voltage delivery were
satisfactory during each of the service propulsion firings. The batteries
were near the fully charged level at entry.
The fuel cells were activated 59 hours prior to launch and the system
was configured with fuel cell 2 on main bus A. Fuel cells 1 and 3 were on
open-circuit until 3.5 hours before lift-off when fuel cells 1 and 2 were
placed on main bus A and fuel cell 3 on main bus B. This configuration was
maintained throughout the flight. Load variance between fuel cells was a
nominal 4 to 7 amperes during flight, with the fuel cells supplying 653
kilowatt-hours of energy at an average current and bus voltage of 77 amperes
and 28.8 volts, respectively.
6.3 CRYOGENIC STORAGE
The cryogenic storage system satisfactorily supplied reactants to
the fuel cells and metabolic oxygen to the environmental control system
throughout the mission. The quantities of oxygen and hydrogen consumed as
compared to preflight predictions are given in section 6.11-3.
The system supplied all demands including the extravehicular activity
during transearth coast when the system supplied a flow rate of
approximately 12.2 lb/hr and the pressure and heater temperatures remained
within the anticipated limits.
Performance of the command and service module communications system was
nominal throughout the mission, except that the Command Module Pilot's
lightweight headset microphone was inoperative when the headset was removed
from stowage. Inflight troubleshooting verified that the failure was in the
microphone. Past history shows three microphone amplifier failures out of
approximately 300 units in use. The headset was transferred to the lunar
module and jettisoned; therefore, the failure could not be isolated to a
The instrumentation performed normally with three exceptions.
a. The service module reaction control system quad A fuel manifold
pressure measurement was intermittently noisy (about 4 percent). However,
there were other measurements for determining the manifold pressure.
b. The central timing equipment timer was reset at 97 hours 53 minutes.
A time correction was inserted by up-data link, and the timer continued to
operate properly throughout the flight. The noise susceptibility of the
reset line to the central timing equipment has been evident on other
spacecraft. However, because of the ease of updating, the problem
has not been considered significant enough to justify redesign.
c. The first 20 feet of tape on the data recorder reproducer became
degraded after about 100 dumps. This portion of the tape was not used for
the remainder of the flight. This anomaly is discussed further in section
6.6 GUIDANCE, NAVIGATION, AND CONTROL
Performance of the guidance, navigation, and the primary and backup
control systems was good throughout the flight. The two anomalies
experienced during the mission were minor in nature causing no loss of
system capability. They were excessive attenuation of light through the
scanning telescope, and improper alignment of the roll axis when the gyro
display alignment pushbutton was depressed. Descriptions of the anomalies
and the corrective action being taken are included in sections 14.1.15 and
The primary guidance system satisfactorily monitored the trajectories
during launch and the translunar injection maneuver. The most probable
velocity errors at insertion were minus 1.5, minus 41.5, and minus 10.8 ft/sec
in the X, Y, and Z platform axes, respectively. The errors were determined
from data obtained from several sources: the Saturn guidance system, the
command module guidance system, the Saturn guidance system data modified by
tracking data, and command module platform realignments in earth orbit.
Separation from the S-IVB and the transposition maneuver were nominal.
Daring the docking sequence, the digital autopilot control mode was changed
from "attitude hold" to "free" while a plus-X translation was being commanded
in order to secure a positive capture latch indication. The body rates
induced by contact and the plus-X thrusting were not nulled and resulted in
misalignment angles of minus 1-1.0, plus 2.2, and plus 1.6 degrees in pitch,
yaw and roll at the start of the retract sequence (see fig. 6-1). The resultant
misalignment caused a greater-than-normal structural loading in the docking
interface (see sec. 7.1).
Figure 6-1. -Rate and attitude error data during transposition and docking sequence.
Body rate transients of less than 0.1 deg/sec in all three axes were
caused by jettisoning of the scientific instrument module door and launching
of the subsatellite.
Accelerometer biases and gyro drift terms were stable throughout the
flight. The gyro drift terms were updated only once, at 27:56.
is a summary of preflight histories and inflight performance data of the inertial
is a summary of inertial measurement unit
realignments performed during the mission.
significant control parameters for each of the service propulsion system
During lunar orbital operations between 84 and 95 hours, the command
and service module maintained a period of local horizontal attitude hold
with the scientific instrument module toward the lunar surface. The
evaporator in the primary coolant loop was turned off to prevent
interference with the inflight science activities. The resulting large temperature
oscillations in the coolant loop gave some concern as to how these
temperature excursions would affect the guidance equipment. Since no direct
guidance equipment telemetry measurements of coolant temperatures were
available, an analysis was performed using a thermal model of the coolant
loop and a ground test was performed with non-flight guidance equipment.
Both indicated that considerable temperature attenuation exists in the
coolant loop and that temperatures experienced during the flight were
within acceptable limits. As a result of the thermal analysis, the maximum
temperature limits in the systems operational data book are being raised to
90 F, peak, and 75 F, average, over a 2-hour period. The effect of coolant
temperature oscillations upon accelerometer bias is shown in
Cislunar midcourse navigation exercises were performed during the
transearth phase to again demonstrate the capability to navigate to safe
entry conditions in the absence of communications with earth.
Separation from the service module, the maneuver to entry attitude,
and sensing of 0.05g during entry were all nominal. The command module
dynamics were seen to change suddenly when the parachute failure resulted
in a decrease in lift (sec. 14.1.9).
The guidance system controlled the vehicle attitude and lift vector
during entry and, based on computer readouts, guided the spacecraft to
landing coordinates of 26 degrees 7 minutes 48 seconds north latitude, and
158 degrees 7 minutes 12 seconds west longitude.
Postflight testing of the entry monitor system scroll indicated that
intermittent scribing occurred after drogue deployment. Chemical analysis
revealed an improper mixture of the phenolic resin and the encapsulated
dye which is used to coat the scroll. The trace that was scribed by the
stylus was visible to the crew during entry but was not visible postflight
because the dye and resin did not develop properly. No hardware changes
will be made since only postflight testing of the scroll is affected.
6.7.1 Reaction Control Systems
Performance of the service module reaction control system was normal
throughout the mission except that some service module propellant isolation
valves closed as on previous missions. Indications that valves were closed
were reported by the crew following launch, S-IVB/command and service module
separation, and scientific instrument module door jettison. In all cases,
the valves were recycled open without incident. A more complete discussion
of this anomaly is given in section 14.1.1.
The performance of the command module reaction control system was
nominal throughout the mission. The effects of dumping raw fuel following
the propellant depletion firing sequence and the association of this
procedure with the parachute failure is discussed in section 14.1.9.
6.7.2 Service Propulsion System
Service propulsion system performance was satisfactory during each of
the eight maneuvers. The steady-state pressure data, gaging system data,
and velocity differentials indicated essentially nominal performance.
Engine ignition procedures for lunar orbit insertion and transearth
injection were revised, however, because of a short which developed
in the ignition control circuitry on the downstream side of the bank A
solenoid valve. A discussion of this malfunction is given in section
Previous flight results have shown the inflight mixture ratio to be
significantly less than expected from engine acceptance test data. The
service propulsion system engine was re-orificed to increase the mixture
ratio for this mission.
shows the propellant unbalance for the
two major engine firings compared with the predicted unbalance. The
unbalance at the end of the transearth injection firing was very small
and shows that the modifications to the engine were satisfactory.
6.8 ENVIRONMENTAL CONTROL AND CREW STATION
6.8.1 Environmental Control System
Performance of the environmental control system was satisfactory,
although several discrepancies required corrective action or minor
changes to the planned operations.
Water leakage at the chlorine injection port was noted on two
occasions when the cap was removed for the daily chlorination. Retightening
of the port septum-retention insert by the crew successfully stopped the
leakage (see sec. 14.1.2). The crew also noted the presence of gas in the
water, especially after heavy usage such as at the end of an eating period
(see sec. 14.1.14). Another problem related to drinking water was that,
on two occasions, at 13 1/2 hours and 277 hours, the potable water tank
failed to refill after use while the waste water tank accepted the normal
fuel cell water production. Proper potable tank filling resumed after a
waste water dump at 28 1/2 hours, but the tank failed to refill after meal
preparation at 277 hours (see sec. 14.1-7).
Command module cabin pressure was increased prior to sleep periods
again on this flight to assist in measuring inflight cabin leakage.
Estimates of 0.03 lb/hr during translunar coast and 0.01 lb/hr during
transearth coast were determined from cabin pressure decay data.
Noises were heard from the cabin fans that were believed to have been
caused by an object striking the fan blades. Cycling of the fans allowed
the fans to run normally (see sec. 14.1.13).
Prior to crew transfer for lunar module housekeeping, difficulty was
experienced in obtaining proper lunar module/command module differential
pressure gage readings. The difficulty resulted in insufficient lunar
module pressure decay at cabin pressure equalization. Consequently, extra
lunar module venting was required to obtain additional oxygen enrichment
and assure minimum oxygen concentration for later suited activities.
Radiator outlet temperatures while in translunar coast and lunar orbit were 10 to 150 F higher than preflight thermal studies indicated they would
be. During the flight, calculations using more accurate heat load inputs
resulted in considerably closer predictions, although some degradation of
radiator coating may have contributed to the higher-than-predicted
temperatures. The radiator outlet temperatures were greater than on previous
missions because of the vehicle attitude and higher electrical loads
required to support the scientific instrument module experiments.
During preparations for lunar module jettison, after an apparently
successful hatch integrity check, the differential pressure decreased
between the command module cabin and tunnel, indicating the possibility of a
command module hatch leak. Although a subsequent 10-minute check
demonstrated satisfactory hatch integrity, an inspection of both the lunar
module and command module hatch seals was made. No evidence of contamination
or damage was found. The hatches were reinstalled and a successful hatch
integrity check was performed. The crew had also experienced difficulty in
obtaining an acceptable suit circuit integrity check during the lunar module
jettison preparations. After being unable to pressurize the suit loop more
than 1 psi above cabin pressure, the crew doffed their helmets and gloves,
and the Commander also unzipped his pressure garment assembly, unlocked and
removed the liquid cooling garment connector, and installed a water
connector plug. After rezipping the suit and donning helmets and gloves, a
successful integrity check was completed. Subsequently, because of the hatch
integrity problem previously mentioned, the suit integrity was again broken,
and the suit check had to be repeated. This check was again unsatisfactory
because one suit glove was not properly connected. After making the correct
connection, a final suit circuit test was successfully completed. The delay
resulted in the lunar module being jettisoned one revolution later than
Droplets of water came from two of the three blue (supply) hoses when
they were relocated for the transearth injection firing. Since cabin
humidity continued to be normal and no recurrence of the problem was
observed, most likely the condensation was an effect of the large primary
coolant loop temperature transient on the suit circuit heat exchanger during
During the period of the entry propellant depletion firing, cabin
pressure continued to increase at a rate consistent with the ambient
atmospheric entry pressure. Manual closure of the cabin pressure relief
valves at that time should have prevented any additional inflow. Since use
of onboard gas supplies was not sufficient to account for the change,
apparently the manual valve was not completely closed or abnormally high
leakage occurred. Postflight examination of the cabin pressure relief
valves and the remote operating mechanism was conducted, and no excess
leakage was indicated.
While being used for postflight testing, the side-A shutoff valve on
the main oxygen regulator toggle arm pivot pin was found sheared. No
problem had been reported during the mission. This anomaly is discussed
further in section 14.1.18.
6.8.2 Crew Station/Equipment
The performance of crew equipment was satisfactory. Three items of
equipment were reviewed as a result of problems experienced during the
The command module ultraviolet window filter was inspected to
determine what action may be required to prevent surface scratching and
improve the optical qualities of the filter. A change has been made to the
filter material to improve the abrasive resistance and optical qualities.
Lengthening of the Command Module Pilot restraint tether was
investigated to provide additional reach for the crewman. The current
length is the maximum allowable to preclude loading of the oxygen
The comma nd module crewman optical alignment sight which came loose
from its mount during landing has been inspected. This anomaly and the
corrective action being taken are discussed in section 14.1.19.
6.9 CONTROLS AND DISPLAYS
The controls and displays performed normally with the following
Direct-current bus B and alternating-current bus 2
undervoltage alarms occurred at approximately
33-3/4 hours; subsequently, an integral lighting
circuit breaker was found open. Since the circuits fed by this breaker were
not mission essential, the breaker was not reset. See section 14.1.4 for
further discussion of this anomaly.
6.10 EXTRAVEHICULAR ACTIVITY EQUIPMENT
At approximately 81-1/2 hours, the battery relay bus measurement read
13.66 volts instead of the nominal 32 volts, as evidenced by backup
measurement readings. Movement of the panel 101 systems test meter switch
caused the reading to return to normal. This anomaly is discussed in section
The mission timer on panel 2 stopped at about 125 hours. After
several attempts, the timer was restarted, and it operated properly for the
remainder of the mission. See section 14.1.8 for further discussion of this
During the crew debriefing, the Command Module Pilot stated that the
seconds digit of the digital event timer located on panel 1 became
obscured by a powder-like substance that formed on the inside of the
glass. For further discussion, see section 14.1.11.
Another problem noted during postflight testing of the vehicle was
that the battery charger main A circuit breaker on panel 5 could not be
manually opened. Corrosion was found around the indicator sleeve of the
breaker actuating knob. This anomaly is discussed in section 14.1-17.
The environmental control system and crew equipment performed
successfully throughout the transearth extravehicular activity. Operation of
the new components, including the umbilical, suit control unit, pressure
control valve, oxygen control and communications panels, and the
extravehicular activity warning system was entirely nominal. All checks and
activities went smoothly, and the extravehicular portion lasted less than 40
minutes. Cabin pressure was restored as planned, using the three 1-pound
oxygen bottles from the rapid repressurization system and CMP-flow mode
until 3.0 psia was reached, and then discharging the unused oxygen purge
system to bring the pressure above 5.0 psia. Subsequent depletion of the
residual 2000 psi in the oxygen purge system was accomplished by using it
once to increase cabin pressure prior to a sleep period and on the following
day, when the remainder was allowed to bleed into the cabin at a controlled
The command and service module consumable usage during the Apollo 15
mission was well within the red line limits and, in all systems, was close
to the preflight predicted values.
6.11.1 Service Propulsion Propellant
Service propulsion propellant and helium loadings and consumption
values are listed in the following
The loadings were calculated
from gaging system readings and measured densities prior to lift-off.
6.1-1.2 Reaction Control System Propellant
Service Module.- The propellant utilization and loading data for the
service module reaction control system were as shown in the following
Consumption was calculated from telemetered helium tank pressure
histories and was based on pressure, volume, and temperature relationships.
Command Module.- The loading and utilization of command module reaction
control system propellant were
Consumption was calculated from
pressure, volume, and temperature relationships.
The total cryogenic hydrogen and oxygen quantities available at liftoff and
Consumption values were based on quantity data
transmitted by telemetry.
The water quantities loaded, produced, and expelled during the mission
are shown in the following