The general trajectory profile of this mission was similar to that of previous lunar missions except for a few innovations and refinements in some of the maneuvers. These changes were: (a) The service propulsion system was used to perform the descent orbit insertion maneuver placing the command and service modules in the low-perilune orbit (9.1 miles). (b) A direct rendezvous was performed using the ascent propulsion system to perform the terminal phase initiation maneuver. Tables 6-I and 6-II give the times of major flight events and definitions of the events; tables 6-III and 6-IV contain trajectory parameter information; and table 6-V is a summary of maneuver data.

The launch trajectory is reported in reference 5. The S-IVB was targeted for the translunar injection maneuver to achieve a 2022-mile pericynthion free-return trajectory. The command and service module/ lunar module trajectory was altered 28 hours later by the first midcourse correction which placed the combined spacecraft on a hybrid trajectory with a pericynthion of 67.0 miles. A second midcourse correction, 46 hours later, lowered the pericynthion to 60.7 miles.

After spacecraft separation, the S-IVB performed a programmed propellant dump and two attitude maneuvers that directed the vehicle to a lunar impact. The impact coordinates were 8 degrees 05 minutes 35 seconds south latitude and 26 degrees 01 minute 23 seconds west longitude; 156 miles from the prelaunch target point but within the nominal impact zone.

6.2 LUNAR ORBIT

6.2.1 Orbital Trajectory

The service propulsion system was used to perform the lunar orbit insertion maneuver. The orbit achieved had an apocynthion of 169 miles and a pericynthion of 58.1 miles. After two lunar revolutions, the service propulsion system was again used, this time to perform the descent orbit insertion maneuver which placed the combined spacecraft in an orbit with a pericynthion of 9.1 miles. On previous missions, the lunar module descent propulsion system was used to perform this maneuver. The use of the service propulsion system allows the lunar module to maintain a higher descent propulsion system propellant margin. Both vehicles remained in the low-pericynthion orbit until shortly after lunar module separation. After separation, the pericynthion of the command and service modules was increased to 56 miles and a plane-change maneuver was later executed to establish the proper conditions for rendezvous.

6.2.2 Lunar Descent

Preparations for lunar descent.- The powered descent and lunar landing were similar to those of previous missions. However, the navigation performed in preparation for powered descent was more accurate because of the command and service modules being in the 58.8- by 9.1-mile descent orbit for 22 hours prior to powered descent initiation. While in this orbit, the Network obtained long periods of radar tracking of the unperturbed spacecraft from which a more accurate spacecraft state vector was determined. The position of the command module relative to a known landmark near the landing site was accurately determined from sextant marks taken on the landmark. Corrections for known offset angles between the landmark and the landing site were used to compute a vector to the landing site. This vector was sent to the lunar module. Also, the Mission Control Center propagated this vector forward to the time of landing to predict errors due to navigation. This procedure was performed during the two revolutions before powered descent and a final landing site update of 2800 feet was computed and relayed to the crew. After ignition for the powered descent, the crew manually inserted the update into the computer.

Powered descent.- Trajectory control during descent was nominal, and only one target redesignation of 350 feet left (toward the south) was made to take advantage of a smoother landing area. After manual takeover, the crew flew approximately 2000 feet downrange and,300 feet north (fig. 6-1) because the targeted coordinates of the landing site given to the lunar module computer were in error by about 1800 feet.

Coordinates of the landing point are 3 degrees 40 minutes 24 seconds south latitude and 17 degrees 27 minutes 55 seconds west longitude, which is 55 feet north and 165 feet east of the prelaunch landing site (fig. 6-2). (Further discussion of the descent is contained in section 8.6.)

Lift-off from the lunar surface occurred at 141:45:40, during the 31st lunar revolution of the command and service modules. After 432.1 seconds of firing time, the ascent engine was automatically shut down with velocity residuals of minus 0.8, plus 0.3, and plus 0.5 ft/sec in the X, Y, and Z axes, respectively. These were trimmed to minus 0.1, minus 0.5, and plus 0.5 ft/sec in the X, Y, and Z axes, respectively. Comparison of the primary guidance, abort guidance, and the powered flight processor data shoved good agreement throughout the ascent as can be seen in the following table of insertion parameters. To accomplish a direct rendezvous with the command module, a reaction control system vernier adjustment maneuver of 10.3 ft/sec was performed approximately 4 minutes after ascent engine cutoff. The maneuver was necessary because the lunar module ascent program is targeted to achieve an insertion velocity and not a specific position vector. Direct rendezvous was nominal and docking occurred 1 hour 47 minutes 10 seconds after lunar lift-off.

The lunar module rendezvous navigation was accomplished throughout the rendezvous phase and all solutions agreed closely with the ground solution. The command module which was performing backup rendezvous navigation was not able to obtain acceptable VHF ranging data until after the terminal phase initiation maneuver. The VHF anomaly is discussed in section 14.1.4. Figure 14-7 is a comparison of the relative range as measured by lunar module rendezvous radar and command module VHF, and determined from command module state vectors and the best-estimate trajectory propagations. The VHF mark taken at 142:05:15 and incorporated into the command module computer's state vector for the lunar module caused an 8.8-mile relative range error.

Several sextant marks were taken after this error was introduced. Because the computer weighs the VHF marks more heavily than the sextant marks, the additional sextant marks did not reduce the error significantly. The ranging problem apparently cleared up after the terminal phase initiation maneuver and the VHF was used satisfactorily for the midcourse corrections. Table 6-VI provides a summary of the rendezvous maneuver solutions.

Two hours after docking, the command and service modules and lunar module were oriented to the lunar module deorbit attitude, undocked, and the command and service modules then separated from the lunar module. The lunar module was deorbited on this mission, similar to Apollo 12. The deorbit was performed to eliminate the lunar module as an orbital debris hazard for future missions and to provide an impact that could be used as a calibrated impulse for the seismographic equipment. The reaction control system of the lunar module was used to perform the 75second deorbit firing 1 hour 24 minutes 19.9 seconds after the command and service modules had separated from the lunar module. The lunar module impacted the lunar surface at 3 degrees 25 minutes 12 seconds south latitude and 19 degrees 40 minutes 1 second west longitude with a velocity of about 5500 feet per second. This point was 36 miles from the Apollo 14 landing site, 62 miles from the Apollo 12 landing site, and 7 miles from the prelaunch target point.

6.3 TRANSEARTH AND ENTRY TRAJECTORIES

A nominal transearth injection maneuver was performed at about 148 hours 36 minutes. Seventeen hours after transearth injection, the third and final midcourse correction was performed.

Fifteen minutes prior to entering the earth's atmosphere, the command module was separated from the service module. The command module was then oriented to blunt-end-forward for earth entry. Entry was nominal and the spacecraft landed in the Pacific Ocean less than one mile from the prelaunch target point.

6.4 SERVICE MODULE ENTRY

The service module should have entered the earth's atmosphere and its debris landed in the Pacific Ocean approximately 650 miles southwest of the command module landing point. No radar coverage was planned nor were there any sightings reported for confirmation.