6.0 TRAJECTORY

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.

TABLE 6-I - SEQUENCE OF EVENTS

EVENT
(See Table 6-II for event definitions)
ELAPSED TIME,
hr:min:sec
Range zero
 21:03:02 G.m.t., January 31, 1971
 
Lift-off -
 21:03:02.6 G.m.t., January 31, 1971
 
Translunar injection maneuver,
 Firing time = 350.8 sec
02:28:32
Translunar injection 02:34:32
S-IVB/command module separation 03:02:29
Translunar docking 04:56:56
Spacecraft ejection 05:47:14
First midcourse correction,
 Firing time = 10.1 sec
30:36:08
Second midcourse correction,
 Firing time = 0.65 sec
76:58:12
Lunar orbit insertion,
 Firing time = 370.8 sec
81:56:41
S-IVB lunar impact 82:37:52
Descent orbit insertion,
 Firing time 20.8 sec
86:10:53
Lunar module
undocking and separation
103:47:42
Circularization maneuver,
 Firing time 4 sec
105:11:46
Powered descent initiation,
 Firing time = 764.6 sec
108:02:27
Lunar landing 108:15:09
Start first extravehicular activity 113:39:11
First data from
Apollo lunar surface experiment package
116:47:58
Plane change,
 Firing time = 18.5 sec
117:29:33
Complete first extravehicular activity 118:27:01
Start second extravehicular activity 131:08:13
End second extravehicular activity 135:42:54
Lunar lift-off,
 Firing time = 432.1 sec
141:45:40
Vernier adjustment maneuver,
 Firing time 12.1 sec
141:56:49
Terminal phase initiation 142:30:51
Terminal phase finalization 143:13:29
Docking 143:32:51
Lunar module jettison 145:44:58
Separation maneuver 145:49:43
Lunar module deorbit maneuver,
 Firing time = 76.2 sec
147:14:17
Lunar module lunar impact 147:42:23
Transearth injection,
 Firing time = 149.2 sec
148:36:02
Third midcourse correction,
 Firing time = 3.0 sec
165:34:57
Command module/
service module separation
215:32:42
Entry interface 215:47:45
Begin blackout 215:48:02
End blackout 215:51:19
Drogue deployment 215:56:08
Landing 216:01:58



TABLE 6-II - DEFINITION OF EVENT TIMES

EVENT DEFINITION
Range zero Final integral second before lift-off
Lift-off

Instrumentation unit umbilical disconnect
Translunar injection maneuver Start tank discharge valve opening,
allowing fuel to be pumped to the S-IVB engine
S-IVB/command module separation,
translunar docking, spacecraft ejection,
lunar module undocking and separation,
docking, and command module landing
The time of the event based on analysis of
spacecraft rate and accelerometer data
Command and service module
and lunar module
computer-controlled maneuvers
The time the computer commands
the engine on and off
Command and service module
and lunar module
non-computer-controlled maneuvers
Engine ignition as indicated by the appropriate
engine bilevel telemetry measurement
S-IVB lunar impact Loss of S-band transponder signal
Lunar module descent engine
cutoff time
Engine cutoff established by the beginning of
drop in thrust chamber pressure
Lunar module impact The time the final data point is transmitted
from the vehicle telemetry system
Lunar landing First contact of a lunar module landing pad
with the lunar surface as derived from
analysis of spacecraft rate data
Beginning of extravehicular activity The time cabin pressure reaches 3 psia
during depressurization
End of extravehicular activity The time cabin pressure reaches 3 psia
during repressurization
Apollo lunar surface
experiment package first data
Receipt of first data considered to be valid from
the Apollo lunar surface experiment package
telemetry system
Command module/
service module separation
Separation indicated by
command module/service module
separation relays A and B
via the telemetry system
Entry interface The time the command module reaches
400 000 feet geodetic altitude as indicated
by the best estimate of the trajectory
Begin and end blackout S-band communication loss due to
air ionization during entry
Drogue deployment Deployment indicated by
drogue deploy relays A and B
via the telemetry system
Earth landing The time the command module
touches the water as
determined from accelerometers



TABLE 6-III - TRAJECTORY PARAMETERS

Event Reference
body
Time,
hr:min:sec
Latitude,
deg
Longitude,
deg
Altitude,
mile
Space-fixed
velocity,
ft/sec
Space-fixed
flight-path,
deg
Space-fixed
heading angle,
deg
Translumer phase
Translumar injection Earth 02:34:31.9 19.53 S 141.72 E 179.1 35 514.1 7.48 65.59
Command and service module/
S-IVB separation
Earth 03:02:29.4 19.23 N 153.41 W 4 297.0 24 089.2 46.84
65.41
Docking Earth 04:56:56 30.43 N 137.99 W 20 603.4 13 204.1 66.31 84.71
Command and service module/
lunar module ejection from S-IVB
Earth 05:47:14.4 30.91 N 144.74 W 26 299.6 11 723.5 68.54 87.76
First midcourse correction
  Ignition
  Cutoff

Earth
Earth

30:36:07.9
30:36:18.1

28.87 N
28.87 N

130.33 W
130.37 W

118 515.0
118 522.1

4 437.9
4 367.2

76.47
76.95

101.98
102.23
Second midcourse correction
  Ignition
  Cutoff

Moon
Moon

76:58:12.0
76:58:12.6

0.56 N
0.56 N

61.40 W
61.40 W

11 900.3
11 899.7

3 711.4
3 713.1

-80.1
-80.1

295.57
295.65
Lunar orbit phase
Lunar orbit insertion
  Ignition
  Cutoff

Moon
Moon

81:56:40.7
82:02:51.5

2.83 N
0.10 N

174.81W
161.58 E

87.4
64.2

8 061.4
5 458.5

-9.97
1.3

257.31
338.18
S-IVB impact Moon 82:37:52.2            
Descent orbit insertion
  Ignition
  Cutoff

Moon
Moon

86:10:53.0
86:11:13.8

6.58 N
6.29 N

173.60 W
174.65 W

59.2
59.0

5 484.8
5 279.5

-0.08
-0.03

247.44
246.94
Command and service module/
lunar module separation
Moon 103:47:41.6 12.65 S 87.76 E 30.5 5 435.8 -1.52 241.64
Command and service module
circularization
  Ignition
  Cutoff


Moon
Moon


105:11:46.1
105:11:50.1


7.05 N
7.04 N


178.56 E
178.35 E


60.5
60.3


5 271.3
5 342.1


-0.1
0.22


248.58
248.36
Powered descent initiation Moon 108:02:26.5 7.38 S 1.57 W 7.8 5 565.6 0.08 290.84
Landing Moon 108:15:09.3            
Command and service module
plane change
  Ignition
  Cutoff


Moon
Moon


117:29:33.1
117:29:51.6


10.63 S
10.78 S


96.31 E
95.40 E


62.1
62.1


5 333.1
5 333.3


-0.04
0.01


237.61
241.79
Ascent Moon 141:45:40.0            
Vernier adjustment Moon 141:56:49.4 0.5 N 37.1 W 11.1 5 548.5 0.52 282.1
Terminal phase initiation Moon 142:30:51.1 11.1 N 149.6 W 44.8 5 396.6 0.73 265.0
Terminal phase final Moon 143:13:29.1 11.3 S 76.7 E 58.8 5 365.5 -0.002 265.5
Docking Moon 143:32:50.5 10.18 S 161.87 W 58.6 5 353.5 0.11 268.06
Lunar module jettison Moon 145:44:58.0 3.21 S 21.80 W 59.9 5 344.6 0.133 281.9
Command and service module
separation
Moon 145:49:42.5 0.62 N 39.58 W 60.6 5 341.7 0.119 282.3
Lunar module
ascent stage deorbit
  Ignition
  Cutoff


Moon
Moon


147:14:16.9
147:15:33.1


11.92 S
12.12 S


67.43 E
63.53 E


57.2
57.2


5 358.7
5 177.0


0.018
0.019


267.3
267.7
Lunar module
ascent stage impact
Moon 147:42:23.4 3.42 S 19.67 W 0.0 5 504.9 -3.685 281.7
Transearth injection
  Ignition
  Cutoff

Moon
Moon

148:36:02.3
148:38:31.5

7.41 N
6.64 S

81.55 W
168.85 E

60.9
66.5

5 340.6
8 505.0

-0.17
5.29

260.81
266.89
Transearth coast phase
Third midcourse correction Earth 165:34:56.7 25.77 N 46.43 E 176 713.8 3 593.2 -79.61 124.88
Command module/
service module separation
Earth 215:32:42.2 31.42 S 94.38 E 1 965.0 29 050.8 -36.62 117.11
Entry and landing phases
Entry Earth 215:47:45.3 36.36 S 165.80 E 66.8 36 170.2 -6.37 70.84
Landing Earth 216:01.58.1            



TABLE 6-IV - DEFINITION OF TRAJECTORY AND ORBITAL PARAMETERS

TRAJECTORY PARAMETERS DEFINITION
Geodetic latitude The spherical coordinate measured along a meridian on the earth
from the equator to the point directly beneath the spacecraft, deg
Selenographic latitude The definition is the same as that of the geodetic latitude
except that the reference body is the moon
rather than the earth, deg
Longitude The spherical coordinate, as measured in the equatorial plane,
between the plane of the reference body's prime meridian and
the plane of the spacecraft meridian, deg
Altitude The distance measured along a vector from the center of
the earth to the spacecraft. When the reference body is the moon,
it is the distance measured from the radius of the landing site to the
spacecraft along a vector from the center of the moon to the
spacecraft, ft or miles
Space-fixed velocity Magnitude of the inertial velocity vector referenced to
the body-centered, inertial reference coordinate system, ft/sec
Space-fixed flight-path angle Flight-path angle measured positive upward from the body-centered
local horizontal plane to the inertial velocity vector, deg
Space-fixed heading angle Angle of the projection of the inertial velocity vector onto
the body-centered local horizontal plane, measured positive eastward
from north, deg
Apogee The point of maximum orbital altitude of the spacecraft above
the center of the earth, miles
Perigee The point of minimum orbital altitude of the spacecraft above the
center of the earth, miles
Apocynthion The point of maximum orbital altitude measured from the radius
of the lunar above the moon as landing site, miles
Pericynthion The point of minimum orbital altitude above the moon as measured
from the radius of the lunar landing site, miles
Period Time required for spacecraft to complete 360 degrees
of orbit rotation, min
Inclination The true angle between the spacecraft orbit plane and
the reference body's equatorial plane, deg
Longitude of the ascending node The longitude at which the orbit plane
crosses the reference body's equatorial plane going from
the Southern to the Northern Hemisphere, deg



TABLE 6-V-A - MANEUVER SUMMARY, TRANSLUNAR

Maneuver System Ignition time,
hr:min:sec
Firing time,
sec
Velocity
change,
ft/sec
Resultant pericynthion conditions
Altitude,
miles
Velocity,
ft/sec
Latitude,
deg:min
Longitude,
deg:min
Arrival time,
hr:min:sec
Translunar injection S-IVB 2:28:32.4 350.8 10 366.5 1979 5396 4:14 N 172:24 W 82:15:19
Command and service module/
lunar module separation
from S-IVB
Reaction control 5:47:14.4 6.9 0.8 1980 5550 2:56 N 173:52 W 82:11:20
S-IVB evasive maneuver S-IVB aux. propulsion 6:04:20 80.0 9.5 0 8368 2:05 N 131:52 W 82:01:01
First midcourse correction Service propulsion 30:36:07.9 10.1 71.1 67 8130 2:21 N 167:48 E 82:00:45
Second midcourse correction Service propulsion 76:58:12 0.65 3.5 61 8153 2:12 N 167:41 E 82:40:36



TABLE 6-V-B - MANEUVER SUMMARY, LUNAR ORBIT

Maneuver System Ignition time,
hr:min:sec
Firing time,
sec
Velocity change,
ft/sec
Resultant orbit
Apocynthion,
miles
Foricynthion,
miles
Lunar orbit insertion Service propulsion 81:56:40.7 370.8 3022.4 169.0 58.1
Descent orbit insertion Service propulsion 86:10:53 20.8 205.7 58.8 9.1
Command module/
lunar module separation
Service module
reaction control
103:47:41.6 2.7 0.8 60.2 7.8
Lunar orbit circularization Service propulsion 105:11:46.1 4.0 77.2 63.9 56.0
Powered descent initiation Descent propulsion 108:02:26.5 764.6 6639.1    
Lunar orbit plane change Service propulsion 117:29:33.1 18.5 370.5 62.1 57.7
Lunar orbit insertion Ascent propulsion 141:45:40 432.1 6066.1 51.7 8.5
Vernier adjustment Lunar module
reaction control
141:56:49.4 12.1 10.3 51.2 8.4
Terminal phase initiation Ascent propulsion 142:30:51.1 3.6 88.5 60.1 46.0
Terminal phase finalization Lunar module
reaction control
143:13:29.1 26.7* 32.0* 61.5 58.2
Final separation Service module
reaction control
145:49:42.5 15.8 3.4 63.4 56.8
Lunar moduie de-orbit Lunar module
reaction control
147:14:16.9 76.2 186.1 56.7 -59.8
  *Theoretical values  



TABLE 6-V-C - MANEUVER SUMMARY, TRANSEARTH

Event System Ignition time,
hr:min:sec
Firing time,
sec
Velocity change,
ft/sec
Resultant entry interface condition
Flight-path
angle, deg
Velocity,
ft/sec
Latitude,
deg:min
Longitude,
deg:min
Arrival time,
hr:min:sec
Transearth injection Service propulsion 148:36:02.3 149.2 3460.6 -7.3 36 127 27:02 S 171:30 W 216:26:59
Third midcourse correction Service module
reaction control
165:34:56.7 3.0 0.5 -6.63 36 170 36:30 S 165:15 E  



6.1 LAUNCH AND TRANSLUNAR TRAJECTORIES

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.)

Figure 6-1.- Crossrange and altitude plotted against downrange during the final phase of descent.

Figure 6-2.- Lunar module landing site on lunar topographic photo map of Fra Mauro.



6.2.3 Lunar Ascent and Rendezvous

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.



TABLE 6-VI - RENDEZVOUS SOLUTIONS

Maneuver Computed velocity change, ft/sec
Network Lunar module Command and service module
Terminal phase initiation Vx = 63.0
Vy = 1.0
Vz = 67.0
Vt = 92.0
Vx = 62.1
Vy = 0.1
Vz = 63.1
Vt = 88.5
Vx = -67.4
Vy = 0.5
Vz = -69.2
Vt = 96.6
First midcourse correction No ground solution. Vx = -0.9
Vy = 0.2
Vz = 0.6
Vt = 1.1
Vx = 1.3
Vy = -0.1
Vz = -1.1
Vt = 1.7
Second midcourse correction No ground solution. Vx = -0.1
Vy = 0.1
Vz = -1.4
Vt = 1.6
Vx = 0.6
Vy = -0.2
Vz = -2.2
Vt = 2.3



6.2.4 Lunar Module Deorbit

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.

Chapter 7 - Command and Service Module Performance Table of Contents Apollo 14 Journal Index