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.


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.


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 specific component.


The instrumentation performed normally with three exceptions.


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 14.1.16.

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. Table 6-I is a summary of preflight histories and inflight performance data of the inertial components. Table 6-II is a summary of inertial measurement unit realignments performed during the mission. Table 6-III summarizes significant control parameters for each of the service propulsion system maneuvers.

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 figure 6-2.

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 14.1-3.

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. Figure 6-3 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.


6.8.1 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 planned.

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 lunar orbit.

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 mission.

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 umbilicals.

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.


The controls and displays performed normally with the following exceptions.


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 rate.


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 table. 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 table. 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 as follows. Consumption was calculated from pressure, volume, and temperature relationships.

6.11.3 Cryogenics

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

6.11.4 Water

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