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

Translunar docking loads were higher than those of previous missions because of a pitch misalignment angle of 11 degrees between the command and service module and the lunar module/S-IVB prior to docking probe retraction to the hard-docked configuration. The bending moment during translunar docking was computed to be 425,000 inch-pounds which approaches the design limit of 437,000 inch-pounds.

The sequence films from the onboard camera showed no evidence of large structural oscillations during lunar touchdown, and crew comments agree with this assessment. Landing on the lunar surface occurred with estimated velocities of 6.8 ft/sec in the minus X direction, 1.2 ft/sec in the plus Y direction, and 0.6 ft/sec in the plus Z direction. The descent rate at probe contact was 0.5 ft/sec. Following probe contact, the descent engine was shut down while the footpads were still about 1.6 feet above the surface, resulting in the 6.8 ft/sec velocity at footpad contact. Computer simulations indicate 1.0 inch of stroke in each primary strut except the forward strut, for which a 3.0-inch stroke is estimated. The simulations also indicate that the forward footpad was off the surface in the final rest position. The crew stated that the forward footpad was loose and rotated easily, confirming the computer results.

At touchdown, the lunar module was located partially inside a small crater with the rim of the crater directly underneath the descent engine skirt. The descent engine skirt buckled during landing. This is accounted for in the touchdown dynamic analysis, and was expected as the skirt length had been extended 10 inches over that of previous vehicles. This buckling was noted by the crew and confirmed by photographs of the damaged skirt ( fig. 7-1 ).

The crew reported that there was a gap between the exit plane of the skirt and the lunar surface, indicating that buckling was probably caused by a buildup of pressure inside the nozzle due to proximity to the lunar surface, and not due entirely to contact of the nozzle skirt with the lunar surface. The crew also reported that the buckling seemed to be uniform around the skirt periphery and that the exit plane height above the surface was uniform.

The vehicle contact velocity and attitude data at touchdown show that the landing was very stable in spite of the relatively high lunar surface slope at the landing point. The plus-Z and plus-Y footpads contacted the lunar surface nearly simultaneously, providing a nose-up pitch rate of 17 deg/sec and a roll rate to the left of 15 deg/sec. Final spacecraft settling occurred 1.8 seconds later. The vehicle at-rest attitude, as determined from the gimbal angles, was 6.9 degrees pitch up and 8.6 degrees roll to the left, resulting in a vehicle tilt angle on the lunar surface of approximately 11 degrees from the horizontal ( fig. 7-2).

The performance of the electrical power distribution system and batteries was satisfactory. Descent battery management was performed as planned, all power switchovers were accomplished as required, and parallel operation of the descent and ascent batteries was within acceptable limits. The d-c bus voltage was maintained above 28.9 volts, and the maximum observed current was 74 amperes, during powered descent. Electrical power used during the mission is given in section 7.9.6.


The steerable antenna exhibited random oscillation characteristics identical to those experienced on previous missions, and resulted in three instances of temporary loss of voice and data. Also at random times, small oscillations were noted and were damped out. The problems with the antenna are discussed in section 14.2.4.

The lunar module did not receive VHF transmissions from the command module during the descent phase of the mission. The checklist erroneously configured the command module to transmit on 296.8 MHz and the lunar module to receive on 259.7 MHz.

With the exceptions noted above, all functions including voice, data, and ranging of both the S-band and the VHF equipment operated satisfactorily during all phases of the mission.


The landing radar acquisition of slant range and velocity was normal. The acquired slant range of 42,000 feet increased to about 50,000 feet in approximately 10 seconds. The indication of range increase may have been caused by blockage from a lunar mountain at initial acquisition. Velocity was acquired at an altitude of approximately 39 000 feet above the local terrain. Landing radar outputs were affected at an altitude of about 30 feet by moving dust and debris.

Rendezvous radar tracking operation during the rendezvous sequence was nominal. A lunar module guidance computer program was used after lunar orbit insertion to point the rendezvous radar antenna at the command and service module, thus enabling acquisition at approximately 109 miles. Two problems were noted during the mission and are as follows:


The controls and displays performed normally. The range/range-rate tape meter glass was broken with about 20 percent of the glass missing; however, the meter operated satisfactory during the flight. Section 14.2.8 contains a discussion of this anomaly.


Guidance, navigation, and control system performance was satisfactory throughout the mission except for two anomalies. There was a simultaneous indication of an abort guidance system warning light and master alarm on two occasions (sec. 14.2.6), and no line-of-sight rate information was displayed on the Commander's crosspointers during the rendezvous braking phase (sec. 14.2-7). Neither anomaly affected overall systems performance.

The primary guidance system was activated at 98 hours 26 minutes, the computer timing was synchronized to the command module computer, and the lunar module platform was aligned to the command module platform. The crew had difficulty seeing stars in the alignment optical telescope while performing the docked realignment of the lunar module platform, but this is normal because of reflected light from the command- module structure. Table 7-1 is a summary of all platform realignments of the primary guidance system, both in orbit and on the lunar surface. The calculated gyro drift rates compare well with the 1 sigma value of 2 meru and indicate good gyro performance. Accelerometer performance was equally good although shifts in the X- and Y-accelerometer bias of 0.39 and 0.46 cm/sec2, respectively, were detected while on the lunar surface. The shift resulted from removing and reapplying power to the inertial measurement unit and is not unusual. Table 7-II is a summary of preflight inertial component calibration data.

After a nominal separation from the command module, the abort guidance system was activated, initialized, and aligned to the primary guidance system. Table 7-III is a summary of preflight and inflight performance of the abort guidance system accelerometers and gyros.

The powered descent to the lunar surface was initiated on time. Table 7-IV is a sequence of significant events during descent. A landing site update to move the targeted landing point 853 meters (2800 feet) downrange was made 95 seconds after ignition. The computer began accepting landing radar updates and began adjusting its estimate of altitude upward by 4800 feet. After enabling landing radar updates, the primary guidance altitude flattened out for approximately 70 seconds ( fig. 7-3). This resulted from the initial estimate of the slope stored in the computer being 1 degree; whereas, the true mare slope was zero. Convergence occurred rapidly once the lunar module was over the Apennine foothills where the computer-stored slope agreed more closely with the actual slope. Figure 7-3 is a time history of altitude from the primary and abort guidance systems. Data indicate that 18 separate deflections of the hand controller were made for landing point redesignations during the approach phase program. The total effect of the redesignations moved the landing site coordinates 338 meters (1110 feet) uprange and 409 meters (1341 feet) to the north. Touchdown disturbances were nominal despite the 11-degree slope upon which the landing occurred. Figure 7-4 is a time history of spacecraft rates and attitudes at touchdown.

Performance during ascent was nominal. For the first time, accelerometer biases were updated while on the lunar surface to correct for the small expected shifts experienced when the system was powered down. Since the lunar surface bias determination technique had not been totally proven, only half of the measured shift in the X accelerometer bias was corrected. As a result, some bias error existed during ascent and contributed about 2 ft/sec to the radial velocity error. Analysis is continuing to determine the cause of the remainder of the radial velocity error and possible causes will be discussed in a supplement to this report.

Because the primary guidance system radial velocity was greater than that from the powered flight processor and the abort guidance system, the velocity residuals at engine shutdown were trimmed using the abort guidance system solutions.

After trimming velocity residuals, an abort guidance system warning and master alarm occurred. They were reset by the crew and a computer self- test was performed successfully. System performance was nominal before, during, and after the warnings. See section 14.2.6 for further discussion of this anomaly. No vernier adjust maneuver was required, and the direct rendezvous was accomplished without incident. Table 7-V is a summary of rendezvous maneuver solutions.

The Commander reported that there were no line-of-sight rate data displayed on his crosspointers at a separation distance of 1500 feet. However, line-of-sight rates existed at this time because thrusting upward and to the left was required to null the -line-of-sight rates. Also, the Command Module Pilot verified the presence of line-of-sight rates. Section 14.2.7 contains a discussion of this anomaly.

After a successful docking, the lunar module was configured for the deorbit maneuver and jettisoned. The velocity changes observed at jettison in the X, Y, and Z axes were minus 1.24, minus 0.01, and minus 0.05 ft/sec, respectively. This is equivalent to a 206 lb-sec impulse. For comparison, the separation velocities observed at undocking prior to powered descent were minus 0.18, minus 0.02, and minus 0.04 ft/sec in the X, Y, and Z axes, respectively, or an impulse of 205 lb-see. The close agreement indicates the tunnel was completely vented and the impulse was due entirely to the separation springs. After jettison, the deorbit maneuver was accomplished and performance was nominal.


7.7.1 Reaction Control System

The reaction control system performed satisfactorily throughout the mission with no anomalies. Skillful use of the system by the crew accounted for the propellant consumption being well below predicted levels. Section 7.9.3 contains a summary of the consumables usage during the mission.

7.7.2 Descent Propulsion System

Data analysis indicates that the descent propulsion system performed nominally during powered descent. The total firing time was 739.2 seconds . The propellant quantity gaging system indicated about 1055 Pounds of usable propellant remained at engine shutdown or about 103 seconds of hover time. The supercritical helium system operated nominally. The skirt of the engine was buckled during landing (sec. 7.1). Section 7.9.1 contains a summary of the descent propulsion system consumables usage during the mission.

7.7.3 Ascent Propulsion System

The ascent propulsion system performance during the lunar ascent maneuver and the terminal phase initiation maneuver was satisfactory. The total engine firing time for the two maneuvers was 433.6 seconds. The ascent propulsion system consumables usage is summarized in section 7.9.2.


The performance of the environmental control system was satisfactory throughout the mission. The waste management system functioned as expected; however, the urine receptacle valve was inadvertently left open for about 6 hours during the first lunar sleep period. This resulted in the loss of about 8 pounds of descent stage oxygen before the crew was awakened to close the valve.

The overspeed of the water separator which occurred on Apollo 14 during cabin-mode operation was not evident during this mission because of a decrease in flow with the helmet and gloves off that resulted from a reconfiguration of valves and hose connections. The only off-nominal performance of the water separator occurred following the cabin depressurization for the standup extravehicular activity when the speed decreased, causing a master alarm (see sec. 14.2.2).

After the first extravehicular activity, a broken quick disconnect between the water bacteria filter and the water drink gun resulted in spillage of about 26 pounds of water into the cabin (see sec. 14.2.3). The water was cleaned up by the crew before the second extravehicular activity.

Fluctuations in water/glycol pump differential pressure were noted following the cabin depressurizations for the standup extravehicular activity and the second extravehicular activity (see sec. 14.2.1). Otherwise, the heat transport system functioned normally.

On Apollo 15, the suits were removed and dried for more than 1 hour by connecting the oxygen umbilicals to the suits and allowing gas to flow through them. This was accomplished at the beginning of each rest period following the first two extravehicular activities.


All lunar module consumables remained well within red-line limits.

7.9.1 Descent Propulsion System Propellant.- The descent propulsion system propellant load quantities shown in the following table were calculated from known volumes and weights of offloaded propellants, temperatures, and densities prior to lift-off. (Figure)

Supercritical helium.- The quantities of supercritical helium were determined by computations using pressure measurements and the known volume of the tank. (Figure)

7.9.2 Ascent Propulsion System Propellant.- The ascent propulsion system total propellant usage was approximately as predicted. The loadings shown in the following table were determined from measured densities prior to launch and from weights of off-loaded propellants. (Figure)

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

7.9.3 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. (Figure)

7.9.4 Oxygen The actual quantities of oxygen loaded and consumed are shown in the following table:

7.9.5 Water

The actual water quantities loaded and consumed, shown in the following table , are based on telemetered data.

7.9.6 Electrical Power

The total battery energy usage is given in the following table.