Gentle to moderate southerly winds extended from the surface to an altitude of 25 000 feet at launch time. The maximum wind was 37 knots at 45 000 feet. Broken thin cirrostratus; clouds covered much of the sky at 25 000 feet, but no low or middle clouds were observed. Surface temperature was 83 F and visibility was 10 miles.


The Apollo 15 launch complex was struck by lightning on five different days during the checkout activities. In all, eleven separate strikes were recorded between June 14 and July 21, 1971. The direct damage incurred was limited to the command and service module ground support equipment.

Launch site lightning strikes have not been documented for program other than Apollo and Gemini. Incidents reported for these two programs are as follows:

Existing weather data were examined for the May-through-July periods from 1966 to the launch of Apollo 15. Reported cloud-to-ground lightning strikes for a period of 90 days within the general vicinity of Cape Kennedy showed the daily average to be as follows:
( Figure)

Thunderstorms are more prevalent on the west side of the Indian River and remain west of the launch pad. During the summer of 19T1, however, the west winds prevailed more frequently than the preceding several years, thus causing the thunderstorms to move east. The lightning density in conjunction with the general easterly movement of the storms contributed to the number of strikes being higher than in the past.


The eighth manned Saturn V Apollo space vehicle, AS-510, was launched on an azimuth 90 degrees east of north. A roll maneuver was initiated at 12.2 seconds after lift-off and the vehicle was placed on a flight azimuth of about 80 degrees. The trajectory parameters from launch through translunar injection were nominal. Earth-parking-orbit insertion conditions were achieved 4.4 seconds earlier than planned.

The performance of the S-IC propulsion system was satisfactory and the specific impulse and mixture ratios were near the predicted values. Four of the eight S-IC retromotors and all of the S-II stage ullage motors were removed for this flight; therefore, the S-IC/S-II separation sequence was revised. This sequence change extended the coast period between S-IC outboard engine cutoff and S-II engine start command by one second. The S-IC/S-II separation sequence and S-II engine thrust buildup performance was satisfactory.

The S-II propulsion system performed normally. The specific impulse and mixture ratios were near predicted values. This was the second S-II stage to incorporate a center-engine liquid-oxygen feedline accumulator as a longitudinal oscillation (POGO) suppression device. The operation of the accumulator system was effective in suppressing these types of oscillations.

The S-IVB stage J-2 engine operated satisfactorily throughout the first and second firings and had normal start and cutoff transients. The firing time for the first S-IVB firing was 141.5 seconds, 3.8 seconds less than predicted. Approximately 2.6 seconds of the shorter firing time can be attributed to higher than predicted S-IVB performance. The remainder can be attributed to S-IC and S-II stage performances. The specific impulse and engine mixture ratio were near the predicted values.

Abnormal temperatures were noted in the turbine hot gas system between the first S-IVB firing engine cutoff and second firing engine start command. Most noticeable was the fuel turbine inlet temperature. During liquid hydrogen chilldown, this temperature decreased from +130 to -100 F at the time of the second engine start command. The oxidizer turbine inlet temperature also indicated a small decrease. In addition, the fuel turbine inlet temperature indicated an abnormally fast decrease after engine cutoff for the first firing. A possible cause of the decrease in turbine inlet temperature was a small leak past the gas generator fuel inlet valve.

The S-IVB firing time for translunar injection was 350.8 seconds. Upon completion of the spacecraft separation, transposition, docking, and extraction operations, the S-IVB evasive maneuver was performed and, subsequently, the vehicle was placed on a trajectory to impact the lunar surface in the vicinity of the Apollo 14 landing site. The S-IVB/instrumentation unit impacted the lunar surface at 1 degree 31 minutes south latitude and 11 degrees 49 minutes west longitude with a velocity of 8455 ft/sec. This impact point is approximately 146 kilometers (79 miles) from the target of 3 degrees 39 minutes south latitude and 7 degrees 35 minutes west longitude. Although the impact location was not within the preferred region, scientific data were obtained from the impact.

The impact point projected from the first auxiliary propulsion system maneuver was perturbed by unplanned attitude control thrusting that occurred to counteract forces resulting from a water leak in the sublimator. Following the second auxiliary propulsion system maneuver, the small and gradually decreasing unbalanced force from the sublimator water leak continued to act for a period of 5 hours and further perturbed the point of impact.

The structural loads experienced during the S-IC boost phase were well below design values. Thrust cutoff transients experienced were similar to those of previous flights. During S-IC stage boost, 4- to 5hertz oscillations were detected beginning at approximately 100 seconds. The maximum amplitude measured at the instrumentation unit was ±0.06g. Oscillations in the 4- to 5-hertz range have been observed on previous flights. The structural loads experienced during the S-IVB stage firings were well below design values.

The guidance and navigation system provided satisfactory end conditions for the earth parking orbit and translunar injection. The control system was different from that of Apollo 14 because of redesigned filters and a revised gain schedule. These changes were made to stabilize structural dynamics caused by vehicle mass and structural changes and to improve wind and engine-out characteristics.

The launch vehicle electrical systems and emergency detection system performed satisfactorily throughout all phases of flight. Operation of the batteries, power supplies, inverters, exploding bridge wire firing units, and switch selectors was normal. Vehicle pressure and thermal environments in general were similar to those experienced on earlier flights. The environmental control system performance was satisfactory. All data systems performed satisfactorily through the flight.

More details of the launch vehicle operation and performance are given in reference 1.