3902-H012-RO-000
TRW NOTE NO. 66-FMT-402
P R O J E C T A P O L L O , ] 2 jggg
A P O L L O MISSION A S - 2 0 6 AS P A C E C R A F T R E F E R E N C E T R A J E C T O R Y
Prepared byTrajectory Design Section
TRW Systems
VOLUME ITRAJECTORY DESCRIPTION
28 FEBRUARY 1966
-CR-1332B3) A P O L L C M I S S I O N AS-206AC B A F T B E F E R E B C E T B A J E C 1 C B Y . V O I O M E
1: TRAJEC10BY DESCRIPTION ( T R W SystemsJ G r o u p ) 92 p
N73-73426
Unclas00/99 17962
Prepared forMISSION PLANNING AND ANALYSIS DIVISION
NATIONAL AERONAUTICS AND SPACE ADMINISTRATIONMANNED SPACECRAFT CENTER
HOUSTON, TEXAS
3902-H012-RO-000
RKP
TRW NOTE NO. 66-FMT-402
P R O J E C T APOLLO
A P O L L O MISSION AS-206AS P A C E C R A F T R E F E R E N C E T R A J E C T O R Y
28 FEBRUARY 1966
VOLUME ITRAJECTORY DESCRIPTION
Prepared for . -MISSION PLANNING AND ANALYSIS DIVISION
NATIONAL AERONAUTICS AND SPACE ADMINISTRATIONMANNED SPACECRAFT CENTER
HOUSTON, TEXASNAS 9-4810
Approved by:.Co Ro Huss, ChiefFlight AnalysisBranchNASA/MSC
Approved by:.JohhlP. Mayer, ChiefMissipn Planning andAnalysis DivisionNASA/MSC
Prepared by: R. M. DiamondW. B. GrayR. D. ShinkleF. J. SichiT. L. Vogr~^TRW Syste
Approved by:C. V. Stableford^Acting ManagerMission Design DepartmentTRW Systems
Approved by:. W.^Pittman, Manager
' Mission Planning andOperationsTRW Systems
TRWSYSTEMS
FOREWORD
This report, which defines the Spacecraft Reference
Trajectory for Apollo Mission AS-206A, is submitted to
NASA/Manned Spacecraft Center by TRW Systems in par-
tial response to Task MSC/TRW A-21 (Establishment of the
Reference Trajectory for Apollo Mission AS-206A), with«
Amendment Number 1, of the Apollo Mission Trajectory
Control Program (Contract No. NAS 9-4810).
This report is presented in three volumes. Volume
I summarizes the mission requirements, the unclassified
spacecraft input data used in the generation of the mission
profile, and the significant trajectory analyses. Volume I
also contains graphical and tabular time histories of perti-
nent spacecraft attitude, position, motion, separation char-
acteristics, and MSFN coverage data. Volume II of this
report contains the trajectory listing of selected mission
profile parameters, and the confidential spacecraft input
data specifications. Volume III presents detailed time his-
tories of radar range, range rate, elevation, elevation rate,
azimuth, azimuth rate, and four look-angles for the MSFN
ground stations available for operation during this mission.
111
• CONTENTS
Page
1. INTRODUCTION AND SUMMARY . . ................ 1
1. 1 Purpose ..................... . ......... 1
1. 2 Scope .............................. • • • 1
1. 3 Mission Profile Summary ................... 2
2. SPACECRAFT MISSION .REQUIREMENTS. ... ......... 9
2. 1 Spacecraft Test Objectives ................... 9
2. 1. 1 Primary Spacecraft Test Objectives ....... 92. 1. 2 Secondary Spacecraft Test Objectives ...... 9
2. 2 Mission Profile Guidelines ... ............... 9
2. 2. 1 Launch Vehicle Guidelines ........ ...... 92. 2. 2 MSFN Coverage /Flight Control Guidelines ... 92. 2. 3 LEM Attitude Guidelines . . . . . ..... ..... 102. 2. 4 LEM Propulsion Guidelines ............ ". 112. 2. 5 General Mission Guidelines . . .......... . 12
3. SUMMARY OF INPUT DATA . . . . ........ . . . ..... :. 13
3. 1 Saturn IB Launch Vehicle .................... 13
3. 2 Spacecraft (LEM-1) ................. ...... 13
3. 3 MSFN Stations ........................... .13
3. 4 Earth Constants and Conversion Factors ......... 14
3. 4. 1 Earth Constants ..................... 153. 4. 2 Conversion Factors . ................. 15
3. 5 Spacecraft and Attitude Reference CoordinateSystems ............... . ............... 16
3. 5. 1 Spacecraft Coordinate System,. X«, Y^, Zg. . . 16
3. 5. 2 Launch S_ite Inertial Reference System,x r Y r z i . . . . . . . . . . . . . . • • • . . . . . . ' • • 1 6
3. 5. 3 Launch Site Rotating Reference System,XR' YR' ZR
3. 5. 4 Relative Vehicle Coordinate System,XRV YRV ZRV
CONTENTS (Continued)
Page
4. MISSION DESCRIPTION. 25^ ' f -
4. 1 As cent-to-Orbit . 25
4. 2 S-IVB/SLA/LEM Orbital Coast .'.. . . . . . . . . . .26
4. 3 Spacecraft Separation 26
4. 4 Orbital Cold-Soak to First DPS Burn 27
4. 5 First DPS Burn . . . . 27
4. 6 Orbital Coast to Second DPS Burn. . 28
4. 7 Second DPS Burn/FITH Abort Test/First APS Burn. . 28
4. 7. 1 Second DPS Burn 284. 7. 2 FITH Abort Test/First APS Burn . . . . . . . . . 30
4. 8 Orbital Coast to Second APS Burn 30
4. 9 Second APS Burn . . 30
4. 10 Orbital Coast to Third APS Burn. . 32
4.11 Third APS Burn. 32
4. 12 Orbital Cold-Soak to Fourth APS Burn 32
4. 13 Fourth APS Burn 33
4. 14 Final Orbital Coast 33
4. 15 Spacecraft Orbital Lifetime Estimates 33
5. NOMINAL TRAJECTORY DATA 41
5. 1 Mission Profile Data ' . - . . . . 41
5. 2 Trajectory Phase Data 41
6. MSFN COVERAGE DATA ' 67
7. SUMMARY OF TECHNICAL ACHIEVEMENT 79
REFERENCES 81
VI
TABLES.
Page
3-1 MSFN Ground Stations and Capabilities 17: - '3i • • '. j
4-1 Ballistic Coefficients : 34
4-2 . Spacecraft Orbital Lifetime Estimates. 34
5-1 Time Sequence of Events 42
5-2 Orbital Characteristics of the Spacecraft Coast Phases. . 44
5-3 Spacecraft Earth Shadow Data. . . . . . 45
5-4' • Ascerit-to-Orbit; Discrete Events Summary . . . . . . . . . 46
5-5 SLA Petal Deployment and Spacecraft Separation;Discrete Events Summary 47
; ' . . • " - • - . " : . ' ' ' ' • - '
5-6 First DPS Burn; Discrete Events Summary. . . . . . . . . . 47
5-7 Second DPS Burn/FITH Abort Test/First APS Burn;Discrete Events Summary.. ... . . . . 48
5-8 Second APS Burn; Discrete Events Summary. . . . . ; . .-.' 48
5-9 Third APS Burn; Discrete Events Summary 49
5-10 Fourth APS Burn; Discrete Events Summary. . 49
6-1 MSFN Mission Coverage 68
6-2 Communication Void Intervals 74
Vll
- ILLUSTRATIONS
Page
1-1 Mission Profile Summary '. . ' 5'
3-1 Saturn IB Launch Vehicle Outboard Profile. . . . . . . . . . 18
3-2 Spacecraft Reference Dimensions 19
3-3 LEM-1 Outboard Profile . . . . . . . . . . . . . . . . . . . . . . 20
3-4 Australia Apollo Tracking Ship Placement Study/Spacecraft Separation . . . 21
3-5 Australia Apollo Tracking Ship Placement Study/First DPS Burn 21
3-6 Bermuda Apollo Tracking Ship Placement Study 22
3-7 Spacecraft Coordinate System. 23
3-8 Spacecraft Reference Coordinate System 23
3-9 Relative Vehicle Coordinate System 24
4-1 Spacecraft Separation; Relative Velocity andDistance during RCS Thrusting 35
4-2 Spacecraft Separation; Relative Velocity andDistance History for Approximately 100 Minutes . . . . . . 35
4-3 Spacecraft Cold-Soak Attitude Orientation. . . . . . . . . . . 36
4-4 Cold-Soak Attitude Orientation History for Launches,Second Quarter 1967. 37
4-5 First DPS Burn Thrust Profile 38
4-6 Second DPS Burn Thrust Profile 38
4-7 Effect of Assumed IMU Drift on Orbit DimensionsFollowing Second DPS Burn , 39
4-8 FITH Staging; Relative Velocity and DistanceHistory During First APS Burn 39
4-9 FITH Staging; Relative Velocity and DistanceHistory" to 100 Seconds 40
4-10 Effect of Assumed IMU Drift on Orbit DimensionsFollowing Second APS Burn 40
ix
ILLUSTRATIONS (Continued)
Page
5-1 Earth Ground Track; Revolutions One and Two . 50
5-2 Earth Ground Track; Revolution Three 51
5-3 . Earth Ground Track; Revolution Four 52
5-4 Earth Ground Track; Revolutions Five through Nine. ... 53
5-5 As cent-to-Orbit; Altitude, Geodetic Latitude,and Longitude . . . 54
5-6 As cent-to-Orbit; Inertia! Velocity, Inertia!Flight Path Angle, and Inertia! Azimuth . . . 54
5-7 ' As cent-to-Orbit; Dynamic Pressure and MachNumber 55
5-8 As cent-to-Orbit; Launch Site Inertia! AttitudeHistory 55
5-9 Spacecraft Separation; Altitude, Geodetic Latitude,and Longitude 56
5-10 Spacecraft Separation; Inertia! Velocity, Inertia!Flight Path'Angle, and Inertia! Azimuth 56
5-11 Spacecraft Separation; AV History. 57
5-12 First DPS Burn; Altitude, Geodetic Latitude,and Longitude 57
5-13 First DPS Burn; Inertia! Velocity, InertialFlight Path Angle, and Inertial Azimuth. . 58
5-14 First DPS Burn; Launch Site Inertial AttitudeHistory 58
5-15 First DPS Burn; AV History 59
5-16 Second DPS Burn/FITH Staging/First APS Burn;Altitude, Geodetic Latitude, and Longitude 59
5-17 Second DPS Burn/FITH Staging/First APS Burn;Inertial Velocity, Inertial Flight Path Angle, andInertial Azimuth 60
5-18 Second DPS Burn/FITH Staging/First APS Burn;Launch Site Inertial Attitude History £,Q
ILLUSTRATIONS (Continued)
5-19
5-20
5-21 ..
5-22
5-23
5-24
5-25
5-26
5-27
5-28
5-29
6-1
Second DPS Burn/FITH Staging/First APS Burn;AV History
.. - -. • - • i- .Second APS Burn; Altitude, Geodetic Latitude,and Longitude i
Second APS. Burn; Inertial Velocity, Inertia!Flight Path Angle, and Inertial Azimuth
Second APS Burn; Launch Site InertialAttitude History.
Second APS Burn; AV History. . . . . ; . . . . . • . . . . . . .
Third APS Burn; Altitude, Geodetic Latitude," andLongitude
Third APS Burn; Inertial Velocity, Inertial FlightPath Angle, and Inertial Azimuth
Third APS Burn;" AV History . . . . . . . . . . ... ; . . . .
Fourth APS Burn; Altitude, Geodetic Latitude,and Longitude . '-'. . •* . . . . . . .
Fourth APS Burn; Inertial Velocity, InertialFlight Path Angle, and Inertial : Azimuth. . . . . . .'.•". . .
Fourth APS Burn; AV History . . . . . . . . . ; . ' . .
MSFN Coverage Summary
Page
61
61
62
62
63
63
• • • : 64
,- .-•• 64
, • 65
. 65 '
• - 66
76
XI
.. • ABBREVIATIONS
APS Ascent Propulsion System
AGS Abort Guidance Subsystem -., '.
GSM Command and Service Module
DPS Descent Propulsion System
EST Eastern Standard Time
FIT-H Fire-In-The-Hole
GMT Greenwich Mean Time
IMU Inertial Measurement Unit
LEM Lunar Excursion Module
LES Launch Escape Subsystem
LMP LEM Mission Programmer
NASA National Aeronautics and Space Administration
MAX Maximum
MCC Mission Control Center
MIN Minimum
MSC Manned Spacecraft Center
MSFC Marshall Space Flight Center
MSFN Manned Space Flight Network
PCM Pulse Code Modulation
PNGCS Primary Navigation and Guidance Control System
RCS Reaction Control System
SLA Spacecraft LEM Adapter
SLR Radar Slant Range
UHF Ultra High Frequency
VHF Very High Frequency
XI11
ABBREVIATIONS (Continued)
deg Degrees
er Earth Equatorial Radius
ft Feet
hr Hours
km Kilometers
lb Pounds
min Minutes
n mi Nautical Miles
rad Radians
sec Seconds
xiv
1. INTRODUCTION AND SUMMARY
1. 1 PURPOSE
The spacecraft reference trajectory defined in this report is de-
signed for the unmanned AS-206A Mission. The mission profile has
been designed such that the spacecraft mission objectives (Reference 1)
are satisfied without violating the spacecraft constraints or the mission
guidelines. The purpose of this report is to incorporate the mission
design refinements issued since the publication of the Apollo Mission
AS-206A Spacecraft Preliminary Reference Trajectory (Reference 2).
1. 2 SCOPE
The Apollo Mission AS-206A reference trajectory reflects the
mission refinements issued since the publication of Reference 2, in- "
corporates the known systems test requirements, and uses the current
guidance logic proposed for thrust vector and attitude control for the lunar
excursion module (LEM) powered flight maneuvers. The data presented
for the ascent-to-orbit phase are based on the AS-206A launch vehicle
trajectory listing transmitted to MSC by the Marshall Space Flight
Center (MSFC) through the Guidance and Performance Subpanel (Re-
ference 3).
This report consists of three volumes. Volume I presents the
mission requirements, a summary of the unclassified spacecraft input
data used in the generation of the mission profile, a description of the
major phases of the mission, and the trajectory analyses of pertinent
phases. Volume I also contains graphical and tabular time histories
of pertinent spacecraft attitude, position, motion, separation charac-
teristics, and MSFN coverage data.
Volume II of this report contains the classified trajectory listing
of the mission profile and the classified spacecraft input data specifi-
cations.
Volume III presents detailed time histories of radar range, range
rate, elevation, elevation rate, azimuth, azimuth rate, and four look-
angles for the MSFN ground stations available for operation during this
mission. . .
1
1.3 MISSION'^PROFIL'E SUMMARY
The AS-206A mission will be the first launch of a LEM spacecraft.
For mission simulation purposes, launch is assumed to occur at 13:00
hours Greenwich mean time (GMT), 1 April 1967, from launch complex
37B of the Kennedy Spaceflight Center.
Major mission events are illustrated in Figure 1-1. The mission
has been divided into 14 phases for discussion purposes:
1) Ascent-to-orbit .
2) S-IVB/space craft LEM adapter (SLA)/L,EM orbital coast
3) Spacecraft separation
4) Orbital cold-soak to first descent propulsion system(DPS) burn
5) First DPS burn
6) Orbital coast to second DPS burn
7) Second DPS burn/fire-in-the-hole (FITH) abort test/first ascent propulsion system (APS) burn
*8) Orbital coast to second APS burn ,
9) Second APS burn -
10) Orbital coast to third APS burn
11) Third APS burn
12) Orbital cold-soak to fourth APS burn • \
'13) Fourth APS burn
14) Final orbital coast
The Saturn IB launch phase includes the thrusting phases of the
S-IB and S-IVB stages. The boilerplate command and service module
(CSM) is jettisoned during S-IVB thrusting by the launch escape subsystem
(L.ES). S-IVB thrust termination occurs at a radius vector magnitude
of approximately 21, 440, 000 feet (perigee altitude of approximately 85
nautical miles) and a zero-degree inertial flight path angle. The velocity
at insertion is such that the radius vector magnitude at apogee is approxi-
mately 21, 650, 000 feet (altitude of approximately. 120 nautic'alT-'rniles). , ;.
These data correspond to the insertion state vector furnished to MSC by
MSFC through the Guidance and Performance Subpanel (Reference 3). .
The spacecraft is separated from the S-IVB/SLA combination during
the first revolution by the LEM/reaction control.system (RCS) thrusters..
MSFIST coverage of this mission event is provided by the Australia Apollo
tracking ship.
The first DPS burn is performed on the third revolution following
a spacecraft attitude-hold thermal soak (LEM X axis normal to the ecliptic
plane and Z axis toward the sun) of approximately 3 hours. The "Hohmann
descent" guidance philosophy is used for the first DPS burn. During this
burn (approximately 39 seconds), the AV available (approximately 150
feet per second) is used to raise the apogee altitude to approximately 180
nautical miles.
The second DPS burn/FITH abort test/first APS burn maneuver is
performed across the United States during the latter part of the third
revolution and the first part of the fourth revolution. The "lunar landing"
guidance philosophy is used for the second DPS burn (excluding the random
throttling phase). The major portion of the AV available (approximately
7000 feet per second) from this 757-second burn is dissipated out of the
orbit plane. The approximate apogee and perigee altitudes following the
random throttling sequence are 200 nautical miles and 155 nautical miles,
respectively.
FITH staging and the first APS burn are accomplished in a constant
inertia! attitude mode. The spacecraft attitude during the 5-second APS
burn dissipates most of the available AV (approximately 55 feet per second)
out of the orbit plane. The spacecraft orbit following this burn is essen-
tially unchanged from the orbit following the second DPS burn.
The second APS burn is performed over the United States near the
end of the fourth revolution. The "LEM ascent" guidance philosophy is
used for the second APS burn. Most of the available AV (approximately
5200 feet per second) resulting from this 395-second burn is dissipated
out of the orbit plane. At shutdown of the second APS burn, the space-
craft is on an orbit defined by an apogee altitude of approximately 203
nautical miles and a perigee altitude of approximately 163 nautical miles.
Following this mission event, and an attitude-hold orbital coast of approx-
imately 45 minutes, the third APS burn is performed while in sight of the
Carnarvon MSFN station. During the constant-attitude 5-second third
APS burn, the spacecraft thrust vector is oriented such that the orbit
perigee altitude is reduced to approximately 149 nautical miles. Approx-
imately 100 feet per second of AV is available from this burn.
The fourth APS burn is performed during the latter part of the
seventh revolution over the United States following an attitude-hold thermal
soak (LEM X axis normal to the ecliptic plane and Z axis toward the sun)
of approximately 3. 5 hours. During the fourth APS burn (approximately
8 seconds), the spacecraft attitude is maintained in an attitude-hold mode
by the abort guidance subsystem. The AV available from this burn
(approximately 160 feet per second) is used to decrease the spacecraft
perigee altitude to approximately 125 nautical miles.
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SPACECRAFT MISSION' REQUIREMENTS^"' -;
2. 1 SPACECRAFT TEST OBJECTIVES.
The purpose of Apollo mission AS-206A is.to conduct an unmanned
test of an operational L/EM for verification of subsystems operations. The
spacecraft test objectives are summarized in'this section.and presented in
Reference 1. ,
2. 1. 1 Primary Spacecraft Test Objectives . ; ':•
a) Verify operation of the following .L/EM subsystems:primary navigation and guidance control system(PNGCS), RCS, APS, and DPS.
b) Evaluate FITH abort test. • - . - • ' ' ""
2. 1. 2 .Secondary Spacecraft Test Objectives
a) Demonstrate APS operation at low propell ant quantities.
b) Demonstrate operation of the mission programmer.
2 . 2 MISSION PROFILE GUIDELINES - - ' . . . , '/
The mission profile .guidelines have been compiled from References
1, 4 through 8,, and supplemented by data furnished by the Manned Space-
craft Center (MSC) at technical coordination meetings.
2. 2. 1 Launch Vehicle Guidelines .
a) 'The nominal flight, azimuth shall be 72 degrees fromtrue North. • . ... '\ . .
7 •
b) The LES will be utilized to jettison the boilerplateCSM from the S-IVB/SLA/LEM combination whenthe dynamic pressure has decreased to less thanone pound per square foot.
c) The guidance command angular rate limitation shallbe one degree per second for both pitch and yawsteering.
2. 2. 2 MSFN Coverage/Flight Control Guidelines ;
a) Continuous MSFN tracking, telemetry, and groundcommand backup coverage are required from liftoffto S-IVB thrust termination (orbital insertion) plus3 minutes.
b) Continuous telemetry, tracking, and ground command; backup coverage are required for all LEM descent and
ascent stage firings and pre-ignition attitude maneuvers.This coverage should commence at least 2 minutesprior to the programmed ignition and should not endprior to confirmation of shutdown.
c) Telemetry and tracking coverage are required at leastonce per revolution during orbital coasts and thermalsoaks.
d) Telemetry, tracking, and ground command backupcoverage are required at least 3 minutes prior tospacecraft separation. This coverage should continueuntil the LEM can radiate to a MSFN station.
i e) State vector update capability is required near theCarnarvon MSFN station.
f) For mission events requiring command coverage bymore than one MSFN station, it is desired that thecoverage overlap be at least 30 seconds.
g) The FITH abort test/first APS burn shall be performedat an orbital position so that at least three MSFN sta-tions with data record capability can receive data fromthis demonstration.
h) The LEM very high frequency (VHF) slant rangeacquisition limit is approximately 740.nautical miles.
2. 2. 3 LEM Attitude Guidelines
a) Shortly after orbital insertion, the S-IVB/SLA/LEMcombination shall be maneuvered such that the X axislies in the plane of the local horizontal and the minusZ axis is directed downward along the local vertical.This attitude shall be maintained in an orbital rate modeuntil the start of the separation sequence to enhancecommunications.
b) A ±0. 3-degree deadband attitude-hold mode shall bemaintained during the separation sequence.
t c) Shortly after separation from the SLA. the LEM shallbe oriented such that the X axis is aligned to within±15 degrees of a normal to the ecliptic plane and theplus Z axis is directed toward the sun. This attitudeshall be maintained for approximately 3 hours forthermal soak of the DPS injector face.
d) During coast periods when the LEM attitude is nototherwise constrained, the minus Z axis shall be
10
directed downward along the local radius vectorto enhance communications and to provide backupfor VHF pulse code modulation (PCM) telemetry.
e) All orbital coasts requiring specific orientationsshall be maintained in the ±5-degree deadband.attitude-hold mode.
f) . Shortly after the third APS burn, the LEM shall beoriented such that the X axis is aligned to within ±15degrees of a normal to the ecliptic plane and the Zaxis is directed toward the sun for thermal soak ofthe APS injector face and RCS quad III. This attitudeshall be maintained for approximately 3 hours.
2. 2. 4 LEM Propulsion Guidelines
a) LEM separation from the S-IVB/SLA combinationshall be accomplished during the first revolutionusing four RCS thrusters firing for 18 seconds toprovide plus X axis translation. The LEM/SLArestraining, straps shall be severed 2 seconds afterRCS ignition. This RCS firing time provides adequateclearance between the stowed landing gear and theSLA petals should one. of the RCS engines fail to fire.
b) Ullage settling maneuvers shall precede all DPSfirings using four RCS thrusters to provide plus Xaxis translation. These ullage maneuvers shall be6 seconds in duration. . .
.• - _ . • . . •- - - f \- • • -
c) Ullage settling maneuvers shall precede the .second,third, and fourth APS firings using four RCS thrustersto. provide plus X axis translation. These ullagesettling maneuvers .shall be 2. 5 seconds in duration.
d) ' The start sequence of each DPS burn shall be asfollows: ignition and the first 3 seconds of the burnshall be at the 30-percent thrust setting; the next23 seconds shall be at the 10-percent thrust setting.
e) The first DPS burn profile shall consist of the startsequence described in the above paragraph followedby approximately 7 seconds of thrusting at the 92. 5-percent thrust setting. Shutdown of the first DPSburn shall occur approximately 33 seconds afterignition.
f) The second DPS burn profile shall consist of thestart sequence previously described followed by379 seconds at the 92. 5-percent thrust setting; forthe next 300 seconds, the thrust shall decrease.linearly from the 60- to the 10-percent thrust
11
setting. The random throttling portion of the burnshall begin following the thrust decay to the 10-percentthrust setting and shall be 52 seconds in duration. Therandom throttling profile shall consist of 10-secondintervals at the 10-, 50-, 30-, 40-, and 20-percentthrust settings followed by a 2-second burn at the 92. 5-percent thrust setting.
g) The FITH abort test is to be initiated immediately afterthe 92. 5-percent thrust setting of the random throttlingphase.
•h) The first APS burn (initiated with the FITH abort test)shall be approximately 5. seconds in duration.
i) The second APS burn shall be approximately 435seconds induration.
j) The third APS burn shall occur approximately 45minutes after shutdown of the second APS burn.
k) The third and fourth APS burns shall be approximately5 seconds in duration.
1) Demonstration of RCS operation with propellant trans-ferred from the APS tanks shall begin 2 seconds afterthe start of the second APS burn and will continue untilshutdown.
m) The fourth APS burn duration shall deplete the APSpropellants to less than 5 percent of the total available.
2. 2. 5 General Mission Guidelines
a) The predicted orbital lifetime for the spent descentand as.cent" stages should not exceed 3 months.
b) LEM separation from the S-IVB/SL.A shall occurduring the first revolution in the vicinity of theCarnarvon MSFN" station.
c) LEM landing gear deployment, shall occur followingthe orbital cold-soak prior to the first DPS burnwhile the spacecraft is in contact with a MSFN station.
d) Pre-ignition attitude orientation maneuvers shalloccur within MSFN coverage by a station withcommand backup capability.
e) The fourth APS burn will be guided by the abortguidance subsystem (ACS) in the attitude-holdmode.
12
... ,. 3. SUMMARY OF : INPUT DATA . - - : r . ., .
The input data contained in this section include the pertinent MSFN
station specifications and the unclassified quantitative description of the
spacecraft. The spacecraft specifications classified as confidential can
be found in Volume II of this report. These data form the basis for the .
spacecraft reference trajectory in support of Apollo Mission AS-206A.
3. 1 -' SATURN IB LAUNCH VEHICLE "
The Saturn IB launch vehicle which will be, used to insert the S-IVB/
SLA/LEM combination into the initial orbit is comprised of the S-IB and
S-IVB stages. The official launch vehicle data and launch vehicle refer-
ence trajectory. (Reference 9) will be furnished by Marshall Space Flight
Center '(MSFG). The Saturn IB launch vehicle is illustrated in Figure 3-1.
3.2 . SPACECRAFT (LEM-1) . ., .
The spacecraft weight statement was obtained from Reference 8 and
is presented in Volume II of this report. The propulsion characteristics
for both the ascent, and descent stages of the. spacecraft are presented in
Volume II of this report. The spacecraft reference dimensions are pre-
sented in Figure 3-2. The LEM-i outboard profile ,is illustrated in
.Figure 3-3. • • - . - • . .
3 . 3 MSFN STATIONS . . . . , .
The locations and capabilities of the-MSFN stations that are planned
to be available for support of Apollo Mission AS-206A were obtained from
References 5 and 10, respectively. These data are summarized in Table
3-1. The station coordinates given are based on the Fischer ellipsoid
model described by an equatorial earth radius of 6, 378, 166. 0 meters, a
polar earth radius of 6, 356,784. 284 meters, and an earth flattening ratio
of 1/298. 3. MSFN station altitudes are referenced to this ellipsoid and
include any known geoidal separation.
The criteria for selecting the locations of the three Apollo tracking
ships which are listed in Table 3-1 are discussed in the following para-
graphs. The Australia Apollo tracking ship (Ship Number One) is planned
for location off the western coast of Australia. The coverage afforded by
13 .
the Carnarvon tracking station and this tracking ship will be used for the
spacecraft separation sequence, state vector updates, the first DPS burn,
:and the third APS burn. The location of this ship is such that the 30-sec-
ond coverage overlap requirement between this station and Carnarvon and
the 3 minutes of coverage prior to separation requirement are satisfied.
This location also provides --at-least 2 minutes of coverage prior to the
first DPS burn and the third APS burn. Summaries of the study data used
in the selection of the location of this ship are illustrated in Figures 3-4
and 3-5.
The California Apollo tracking ship (Ship Number Two) is to be
located off the western coast of the continental United States. Coverage
provided by this station is required for the second DPS burn and the sec-
ond and fourth APS burns. Because of the large central angle traversed .
by the spacecraft during the second DPS burn, it was required that the
location of this ship be positioned to increase coverage for this mission
event, yet provide the necessary coverage for the second and fourth APS
burns. To satisfy these requirements, this station was located as far off
the western coast of the United States as possible, while still retaining a
30-second coverage overlap with the Texas MSFN station.
The Bermuda Apollo tracking ship (Ship Number Three) is located
so that it provides coverage from 30 seconds prior to S-IVB shutdown
(orbital insertion) to at least 3 minutes after S-IVB shutdown. A summary
of the study.data used in the selection of the position of this ship is pre-
sented in Figure 3-6. The data presented in Figure 3-6 may. also be used
to adjust the position of this ship and to determine the location of any
additional coverage sources which may be required for support of possible
contingency situations.
3. 4 EARTH CONSTANTS AND CONVERSION FACTORS
The earth constants and conversion factors presented on the follow-
ing page were extracted from Reference 4 and have been used in the
generation of the spacecraft reference trajectory.
14
3. 4. 1 Earth .Constants
Rotational rate
Equatorial radius
Average radius
Gravitational parameter (|a )
4. 37526902 x 10"3 rad/min
0. 417807416 x 10~2 deg/sec .
0.729211504 x 10"4 rad/sec
2. 092573819 x 10? ff' . :
2. 0909841 x 107 ft
5. 53039344 x 10"3 er3/min2
11. 46782384 x 103 er,3/day2
3. 986032 x 105km3 /sec2-
1.407653916 x 1016 ft3 /sec2
Coefficients of potential harmonics
J term .(second)
H term (third) ' '
D term (fourth)
Earth flattening
3. 4. 2 Conversion Factors
Kilometers per foot
Kilometers per nautical mile
Feet per nautical mile
Weight-to-mass ratio
Mass-to-weight ratio
Feet per earth equatorialradius
Nautical miles per earthequatorial radius
-3-1.62345 x 10 nd
-0.575 x 10~5 nd
0. 7875 x.10"5 nd
1/298: 3;nd
0.'3048 x 10~3km/ft
. ,1..852 krn/n mi
.6076.115486 -ft/n mi
32.'17404856 Ib/slug
0.031080950 slug/lb
2. 092573819 x 10? f t /er
3443. 93358 n mi/er
15
3. 5 SPACECRAFT AND ATTITUDE REFERENCE COORDINATE '-.SYSTEMS
The spacecraft and attitude reference coordinate systems are illus-
trated in Figures 3-7 through 3-9. The spacecraft attitude is measured
by the pitch, yaw, and roll angles required to rotate from the reference
..system .to the. current spacecraft orientation. . , . . ______
3. 5. 1 Spacecraft Coordinate System, X „, Yg, Z_
The LEM coordinate system (see Figure 3-7) is an orthogonal,
right-handed system coincident with the spacecraft axes. The Xg axis
extends through the upper docking tunnel. The Zg axis extends along the
crew line -of -sight and the Yg axis completes the right-handed system.
3. 5. 2 Launch Site Inertia! Reference System, XT, YT, Z
This orthogonal, right-handed coordinate reference system (see
Figure 3-8) coincides with the launch site at the time of launch. The Xy
axis extends downrange in the direction of the launch azimuth. The Z
axis extends upward along the astronomical vertical and the YT axis
completes the right-handed system.
3. 5. 3 Launch Site Rotating Reference System, XR, Y ', ZR
This coordinate reference system (see Figure 3-8), which rotates
with the earth,is coincident with the XT, YT, Z system at launch.
3. 5. 4 Relative Vehicle Coordinate System, XR-,, YRV>
This coordinate system (see Figure 3-9) is an orthogonal, right-
handed system centered at the primary vehicle. The X-...... axis extends_ • K. V
in the direction of motion. The YRV axis extends upward along the prima-
ry vehicle position vector. The Z_v axis completes the right-handed
system.
16
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ALL VEHICLE STATIONS AND DIMENSIONS ARE IN INCHES
Figure 3-1. Saturn IB Launch Vehicle Outboard Profile
18
.XL = 379.8;.X =1463.3a
LAUNCHESCAPE-SYSTEM
X = 1000.0BOILERPLATE v -0
CSM X c - °
X =838.0 —a
SPACECRAFT 336.0 IN. / \LEM ADAPTER
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— 154,0 IN DIAMETER
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Figure 3-2. Spacecraft Reference Dimensions
19
S-BAND STEERABLE. . ANTENNA
UPPER DOCKINGTUNNEL
ASCENT STAGE
RENDEZVOUSRADARANTENNA
S-BAND INFLIGHT'ANTENNA (2
VHP ANTENNA (2)
RCS THRUSTERASSEMBLY
RCS NOZZLE
Figure 3-3. LEM-1 Outboard Profile
20
§ 200
. S
-' 55
Figure 3-4. Australia Apollo Tracking Ship Placement Study/Spacecraft Separation
TOTAL COVERAGEOF SELECTED SHIPPOSITION
ILATITUDE OF AUSTRALIAAPOLLO TRACKING SHIP(DEC SOUTH)
SELECT D SHIP POSI ION
MINIMUM OVERLAP TIME (30 SECONDS)
92 14 96 98 100 103 104 106
AUSTRALIA SHIP LONGITUDE (DEC EAST)
Figure 3-5. Australia Apollo Tracking Ship Placement Study/First DPS Burn
21
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Figure 3-8. Reference Coordinate System
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24
4. MISSION DESCRIPTION
The spacecraft reference trajectory for Apollo Mission AS-206A ,
has been designed to satisfy the mission test objectives, (Reference 1) .
and known test requirements without-violating the mission profile .guide-i
lines (Section 2. 2). This mission profile- uses the current L/EM guidance
philosophy for thrust vector and attitude control during the spacecraft
first DPS, second DPS.and second APS burns. The spacecraft first,
third, and fourth APS burns are accomplished in an attitude-hold mode.
The description of the AS-206A mission simulation and the results of
pertinent mission planning studies and mission analyses are presented
in this section.
4 . 1 ASCENT-TO-ORBIT : . . . ; . '
Launch of Apollo Mission AS-206A is planned to occur from launch
complex 37B of the Kennedy Spaceflight Center. ; F'br mission simulation
purposes, launch is assumed to occur at 13:00 hours GMT (08:00 hours
EST) on 1 April 1967. " . - •• '
The ascent-to-orbit profile presented in this report is based upon
an MSEC trajectory listing transmitted to MSC through the Guidance and.
Performance Subpanel (Reference 3). The .ascent-to-orbit phase of the
mission profile utilizes a 72-degree flight azimuth from true North and
includes the burn of the S-IB stage and the burn of the S-IVB stage. The
boilerplate C.SM is jettisoned when the dynamic pressure decreases to
less than one pound per square foot. . " . . ' • :
Insertion into an elliptical earth orbit occurs 603.44 seconds after
liftoff. The state vector at insertion is described as follows:
• Inertial velocity magnitude of 25, 658 feet per second
• Radius vector magnitude of 2.1, 440, 000 feet(approximately 85 nautical miles)
Inertial flight path angle of 0. 0 degree
Orbital plane inclination-of 31. 6.1 degrees. -
25
The ascent-to-orbit phase presented in this report has been de-
signed to satisfy the previously mentioned conditions. MSFN coverage is
provided by the Cape Kennedy, Grand Bahama, Grand Turk, Bermuda, and
the Bermuda Apollo tracking ship stations. It should be noted that the data
for this mission phase are presented for completeness only; the official
launch vehicle^ data will be published by MSFC in the launch vehicle ref-
erence trajectory (Reference 9).
4. 2 S-IVB/SLA/LEM ORBITAL COAST
Shortly after insertion into the initial orbit, the S-IVB/SLA/LEM
combination is maneuvered so that the spacecraft X axis lies in the plane
of the local horizontal and the minus Z axis is directed along the radius,
vector toward the earth. This attitude is maintained by the S-IVB in an
orbital rate mode until 120 seconds prior to LEM separation to enhance
LEM/MSFN communications. SLA petal deployment occurs during the
first revolution (30 minutes after .orbital insertion).
4.3 SPACECRAFT SEPARATION
LEM separation from the S-IVB/SLA combination occurs near
apogee during the first revolution, approximately 10 minutes after SLA
petal deployment. The selection of the in-orbit position for this mission
event is such that continuous MSFN ground command coverage is provided
by the Australia Apollo tracking ship and the Carnarvon station. Sixty
seconds after acquisition the S-IVB begins an attitude-hold maneuver,
maintaining a constant inertial attitude through the spacecraft separation
sequence. Two minutes after the beginning of the inertial attitude-hold
period, four LEM/RCS thrusters are ignited to provide plus X axis
translation. At RCS ignition, the spacecraft plus X axis is in the orbital
plane approximately 9 degrees above the local horizontal. Two seconds
after RCS ignition, the SLA/LEM restraining straps are severed and the
LEM begins the withdrawal from the SLA, Eighteen seconds after ignition
RCS thrusting is terminated. The separation characteristics (LEM-to-
S-IVB relative velocity and distance) are presented in Figures 4-1 and
4-2. The total coverage provided by the two MSFN stations for this event
is approximately 8 minutes and 40 seconds.
26
4. 4 ORBITAL COLD-SOAK TO FIRST DPS BURN " ' ' : :'
Ten seconds after shutdown of the four RCS engines in the. space-
craft separation sequence, the LEM is commanded to perform an attitude
maneuver to align the X axis normal to the ecliptic plane and to direct
the Z axis toward the sun. The spacecraft maintains this inertial attitude
in the ±5-degree deadband mode for the duration of the cold-soak (ap-
proximately 3 hours). The spacecraft attitude orientation during the
cold-soak period is illustrated in Figure 4-3.
Since the sun's position relative to the spacecraft is dependent on
the launch date and time, the maneuver included in this mission profile is
valid only for the assumed launch date and time (13:00 hours GMT,
1 April 1967). The spacecraft inertial attitude angles which satisfy this
thermal cold-soak requirement are presented in-Figure 4-4 as a function
of launch date and time for the second quarter of 1967.
4. 5 FIRST DPS BURN
Upon acquisition of tracking by the Australian tracking ship on
the third revolution, an'orientation maneuver is initiated to align the
spacecraft axes to the desired orientation for'the first DPS burn (see
Figure 5-14). This maneuver is executed in the 5-degree per second
rate mode. Shortly after the termination of this maneuver, the LEM
landing gear is deployed. Approximately one minute has been provided
prior to RCS ignition for MSFN verification of gear deployment.
The first DPS ignition is preceded by a 6-second RCS plus X axis
translation for ullage settling. The first DPS burn start sequence con-
sists of ignition and the first 3 seconds of burn time at the 30-percent
thrust setting, followed by 23 seconds at the 10-percent thrust setting.
This start sequence is followed by a one-second buildup to, and approxi-
mately 11. 5 seconds at, the 9.2. 5-percent thrust setting (see Figure 4-6).
The "Hohmann descent" LEM guidance philosophy is used to con-"
trol the spacecraft attitude and attitude rates during this burn (Reference
11). This guidance philosophy is characterized by burnout at a zero-
degree flight path angle. : •
27
At shutdown of the first DPS burn, the spacecraft is on'an orbit
defined by a predicted apogee altitude of 180 nautical miles, a predicted
perigee altitude of 119 nautical miles, and an orbital period of 90 minutes.
The orbit is inclined 31. 62 degrees to the equator and the longitude at* •' . " ~ "
burnout (perigee) is 96. 5 degrees east. The continuous MSFN ground
command coverage of the first DPS burn is provided by the Australia
tracking ship and the Carnarvon station. The total coverage time is
approximately 9 minutes and 30 seconds.
* It was determined that the second DPS burn would be the least
sensitive to attitude errors if the line of apsides of the orbit following
the first DPS burn would bisect the central angle traversed by the space-
craft during the second DPS burn. An orbit defined by a 120-nautical
mile perigee and a 180-nautical mile apogee resulted in a first DPS
burn duration which satisfied the mission guidelines and provided con-
tinuous MSFN command coverage during the first and second DPS burns.
4. 6 ORBITAL COAST TO SECOND DPS BURN
Following shutdown of the first DPS burn, the spacecraft begins
an orbital coast of approximately 35 minutes. Shortly after the begin-
ning of the coast, a LEM maneuver to orient the X axis in the plane.ofi •
the local horizontal and the minus Z axis toward the geocenter is per- .
formed at the 5-degree per second attitude-rate mode. The spacecraft
is oriented in this manner to enhance LEM/MSFN communications.
This attitude is maintained in the ±5-degree attitude-hold mode.
4.7 SECOND DPS BURN/FITH ABORT TEST/FIRST APS BURN
4^7. 1 Second DPS Burn
At a mission time of approximately .4 hours and 33 minutes the LEM
begins an orientation maneuver to the desired inertia! attitude (Figure
5,-18) for the second DPS burn. The attitude maneuver is performed in
the 5-degree per second rate mode. Approximately 70 seconds after the
completion of the attitude orientation maneuver, a 6-second RCS maneuver
is initiated to provide plus X axis translation for ullage settling immedi-
ately prior to DPS ignition. The desired thrust profile, for the second
DPS burn (Figure 4-7), consists of the start sequence described in
28
Section 4.'5, followed by a 379-second burn" phase- at the -92. 5-rpercent'
thrust setting and a 300-second linear decay from 'the 60- to the 10- •-" '
percent thrust setting. Upon completion of-the linear decay, a 52-second
random throttling sequence is initiated. This throttling profile consists '
of ten-second intervals at the 10-, 50-, 30-, 40-, and'20-percent thrust
settings, respectively. The random throttling profile ends with a two- •
second burn at the 92. 5-percent thrust setting; This burn profile is
typical of the lunar landing mission maneuvers that may be performed
from the completion of lunar de-orbit to lunar;touchdown.,., , .
The guidance philosophy used to provide control of the second DIPS '
burn thrust vector magnitude and direction utilizes multiple targeting •'
conditions (Reference 11). .Each targeting condition consists of the de?-
sired position, velocity, and acceleration vectors and the burn duration,
for the spacecraft to satisfy these conditions. It should be rioted that
the random throttling sequence, FITH staging, arid the first APS.-burn :.
are performed in an attitude-hold mode. , . . . - , .
The second DPS thrust profile resulting from the guidance com-
mands is compared to the desired profile (from the mission guidelines) ,
in Figure 4-7. • : . . , . . . - ' ; .
At shutdown of the second DPS burn, the LEM is- on an orbit with
a predicted perigee altitude of 155 nautical miles, a predicted apogee
altitude of 200 nautical miles, 'a period of approximately 90. 9 minutes,
and an orbital inclination of approximately 30. 5 degrees. An open loop :
analysis to determine the effects of an assumed IMU drift (Reference 7) on
various spacecraft orbit dimensions following the second DPS burn was
performed. The data, generated from this study-are presented in Figure
4-8 and were used in the selection bf'the desired o.rbit dimensions fol-
lowing, the second DPS burn.
The majority of the AV available from this burn (approximately
7, 000 feet per second) is dissipated out of the orbit plane. Approximately-
one-percent of the useable DPS propel-lant remain after the second DPS
burn. - - . , - . . •
29
MSFN coverage of the second DPS burn is provided by the California ;
Apollo tracking ship, Goldstone, Texas, Cape Kennedy, Guaymas, White ;
Sands, Grand Bahama, Eglin, and Bermuda stations. : . :. ,
4. 7. 2 FITH Abort Test/First APS Burn
The FITH a'bort test is initiated immediately at the completion of
the random throttling phase of the second DPS burn with simultaneous
LEM staging and APS ignition signals. The duration of the first APS
burn is 5. 24 seconds. The in-orbit position of this event has been selec-
ted such that MSFN ground command coverage and data record capability
are furnished by the Cape Kennedy, Grand Turk, Bermuda, Grand Bahama,
and Eglin stations. In the mission simulation, 1, 180 pound-seconds
of impulse were delivered to both the LEM ascent and descent stages at
the APS ignition to simulate APS plume impingement effects. Separation
characteristics for the ascent and spent descent stages are presented in
Figures 4-9 and 4-10. At FITH staging, the spacecraft X axis is approxi-
mately normal to the orbital plane. At shutdown, the ascent stage is oil
an orbit defined by a predicted perigee altitude of 152. 5 nautical miles,
a predicted apogee altitude of 200. 0 nautical miles, a period of approxi-
mately 90. 9 minutes, and an inclination of approximately 30. 5 degrees.
The spent descent stage is left in an orbit with a predicted apogee altitude
of 200. 2 nautical miles, a predicted perigee altitude of 153. 4 nautical
miles, an inclination of approximately 3.0. 5 degrees, and a period of
90. 9 minutes. The maximum estimate of the lifetime of the spent descent
stage is 49 days (see. Table 4-2). _—' '
4. 8 ORBITAL COAST TO SECOND APS BURN
Shortly after the first APS burn, a LEM ascent stage attitude ma-
neuver is initiated to align the X axis in the plane of the local horizontal
and to direct the minus Z axis toward the geocenter. This maneuver is
accomplished in the 5-degree per second maneuver mode. This orienta-
tion is maintained in the ±5-degree attitude-hold mode throughout this
coast (approximately 1 hour and 20 minutes).
4. 9 SECOND APS BURN
Following acquisition by the California tracking ship during the
30
four th revolution, an attitude-maneuver is initiated to orient the LEM'
ascent stage to the desired ignition attitude (see Figure 5-22). This
maneuver is performed in the 5-de'gree pe:r second maneuver mode.' •
Two minutes after the acquisition of tracking, the RCS thrusters are
ignited for 2. 5 .seconds of plus X. axis translation for ullage settling.
Immediately after the 2. 5-second RCS burn, the second APS burn is
ini t iated.
The "LEM ascent" guidance, philosophy is used to control the. ;.
spacecraf t attitude, attitude rates., and. second APS burn duration. „ This -
philosophy is characterized by the use of a desired velocity vector at
burnout and 'an intercept time. The burn duration is varied so that the .
velocity vector requirements, are satisfied at the intercept time. The
philosophy used to target this burn consisted of varying the intercept
parameters such that the desired burn time (approximately 395 seconds) ,
and orbital .conditions were met. .It was required that the second APS-
burn duration be reduced from that specified in Reference 1 (approxi-
mately 435 seconds) in order that sufficient propellants remain for .the.
required third and, fourth APS burns. . . . ; .
At shutdown of this burn, the spacecraft orbit is defined by a pre-
dicted apogee altitude of 203. 1 nautical miles, 'a predicted perigee alti-
t u d e of 163. 3 nautical miles, and a period of approximately 91. 2 minutes.
The orbital plane-is inclined approximately 31.4 degrees to the equator.
The desired orbit dimensions at second APS burn shutdown were chosen
so that the orbital lifetime of the ascent stage could be significantly
reduced during the third and fourth APS burns and to reduce the effects
of the assumed IMU drift (Figure 4-10). Continuous MSFN coverage of
this burn is furnished by the California Apollo tracking ship , the Texas,
White Sands, Guaymas, Goldstone, and Point Arguello stations. The
total coverage time is approximately 13 minutes.
Two s-econds after initiationyof the second APS-burn, a test of RCS
operation is initiated. This test is conducted to determine the operation
of the RCS using propellants fed from the APS propellant tanks. The
RCS is used for stabilization of the ascent stage during APS operation.
This test is continued until shutdown of the second APS burn.
31
As in the second DPS burn, the majority of the AV available during
this burn (approximately 5, 200 feet per second) is.dissipated out of the
orbit plane to maintain orbit dimension control.
4. 10 ORBITAL, COAST TO THIRD APS BURN
Shortly after shutdown of the second APS burn, the LEM is com-
manded to maneuver to the inertial attitude described in Section 4. 8 at
the 5-degree per second attitude maneuver mode. This attitude is main-
tained in the ±5-degree deadband mode until acquisition of tracking by
the Australia tracking ship during the fifth revolution (approximately
53 minutes after shutdown of the second APS burn).
4. 11 THIRD APS BURN
Upon acquisition of tracking by the Australia Apollo tracking ship,
the LEM is commanded to maneuver to the desired ignition attitude for
the third APS burn in the 5-degree per second attitude orientation mode.
Two minutes after acquisition of tracking, a 2. 5-second RCS maneuver
is performed to provide plus X axis translation for ullage settling. Im-
mediately after the ullage maneuver, the third APS burn (5. 24 seconds
in duration) is initiated. A constant inertial attitude was selected for this
burn to decrease perigee altitude (thus orbital lifetime). This burn
results in an orbit with an apogee altitude of approximately 220. 1 nautical
miles, a perigee altitude of approximately 149. 4 nautical miles, a 91. 2-
minute period, and an orbit plane inclination of approximately 3 1. 4 de-
grees. Approximately 10 minutes and 30 seconds of continuous MSFN
coverage is furnished for this burn by the Carnarvon and Australia
tracking ship stations.
4. 12 ORBITAL COLD-SOAK TO FOURTH APS BURN
Shortly after completion of the third APS burn the LEM is maneu-
vered such that the X axis is normal to the ecliptic plane and the Z axis
is directed toward the sun for cold-soak of the APS injector face. This
maneuver is accomplished in the 5-degree per second attitude orientation
mode. This attitude is maintained in the ±5-degree deadband mode for
approximately 3 hours and 30 minutes.
32
4. 13 FOURTH APS BURN ." V
Following acquisition by the California tracking ship during the
seventh revolution, an orientation maneuver is initiated in the 5-degree
per second mode to align the spacecraft axes to the desired ignition
attitude for the fourth APS burn. Two minutes after acquisition, 2. 5
seconds of RCS thrusting is initiated to provide plus X axis translation
for ullage settling. This is followed immediately by the fourth APS burn
of approximately 8.2 seconds which depletes all of the available APS
propellarits with the exception of the one percent flight performance . ;
reserves. • • ' . . .
The inertial attitude held during the APS burn further decreases
the perigee attitude. At APS shutdown, the spent spacecraft is on an
orbit characterized by a predicted perigee altitude of 124. 8 nautical
miles, a predicted apogee altitude of 199. 0 nautical miles, an orbital
period of approximately 90. 4 minutes, and an orbital inclination of
approximately 31. 4 degrees. The estimated lifetime of the spacecraft
in this orbit is approximately 19 days.
4. 14 FINAL ORBITAL COAST : . • .
The majority of the mission objectives are completed at this phase
of the mission; however, additional spacecraft tests may be scheduled
after this time. An additional 3 hours have been added after the com-
pletion of the fourth APS burn to provide data for planning any additional
spacecraft testing.
4. 15 SPACECRAFT ORBITAL LIFETIME ESTIMATES
The estimated orbital lifetimes of the various LEM-1 configurations
are presented in Table 4-2. The configurations which have been analy-
zed are defined below:
Configuration 1
The LEM-1 after separation from the S-IVB/SLA combinationand prior to propulsion tests.
33
Configuration 2
The-spent descent stage after the FITH abort test.
Configuration 3
The spent ascent stage after shutdown of the fourth APS burn.
•. Ballisticrcoefficients, W/C.-.A (weight divided by.the orbital drag
coefficient and the frontal area), were calculated for each of the configu-
rations. An orbital drag coefficient of 2. Owas assumed. Various con-
figuration orientations were analyzed to determine the maximum and .
minimum frontal areas which each configuration could exhibit normal to
the velocity vector. Table 4-1 presents the results of this analysis.
Table 4-1. Ballistic Coefficients
Configuration
12233
Area (ft2)
* 200(Maximum) 200(Minimum) 80(Maximum) 190(Minimum) 125
Weight (Ib) '
33, 4755, 1955, 1956, 0576, 057
W/C-A ( lb / f t 2 )JJ83. 69'12. 9932. 4715.9424. 23
*The frontal area of the LEM does not change appreciably.
These ballistic coefficients, the applicable orbital characteristics,
and information presented in Reference 10 were used to calculate the
orbital lifetime estimates presented in Table 4-2.
Table 4-2. Spacecraft Orbital Lifetime Estimates
Orbital Lifetime (days)C onfigur ation Minimum Maximum
1 4 42 19 493 13 19
34
50
40
u*.UJ
•g 30
t—
—OUJ
S 20
UJ
10
>-I—
o
SPACECRAFT SEPARATION OCCURS52 MIN 5 SECFROM LIFTOFF
-4 0 4 . 8 12 16
TIME FROM SEPARATION (SEC)
•20
Figure 4-1. Spacecraft Separation; Relative Velocity and DistanceDuring RCS Thrusting
RCS SHUTDOWNOCCURS 52 MIN 21 SEC •FROM LIFTOFF
100,000
80,000
Z 60,0002£0
S 40,000
20,000
0 L.
12.0
100
> 80tb
- y 60
40
20
•RCS SHUTDOWN
RELATIVEVELOCITY
0 U»I
///
^.RELATIVE./DISTANCE
0:00 0:20 0:40 1:00 1:20 1:40
TIME FROM RCS SHUTDOWN(HR:MIN)
Figure 4-2. Spacecraft Separation; Relative Velocity and DistanceHistory for Approximately 100 Minutes
35
36
. 240
15APR .1MAY / • I5MAY . '.' IJUN
• • ..'; ' '. ' • LAUNCH DATE (1967) "
15JUN I JUL
1APR 15APR 1MAY 15MAY IJUN
LAUNCH DATE (1967)
15JUN I JUL
J
OS
Ou
GREENWICH MEAN TIME AT LAUNCH (HRS)
IAP8 I5APR IMAY I5MAY IJUN
LAUNCH DATE (1967)
I5JUN I JUL
Figure 4-4. Cold -Soak. Attitude Orientation Historyfor Launches, Second Quarter, 1967
37
100
80
60
40
20
\DESIRED THRUSTGUIDANCE COMMANDED THRUST \
0 . 4 6 8 10
TIME FROM DPS IGNITION (MIN)
12 14
Figure 4-5. First DPS Burn Thrust Profile
100
80
1 60
40
20
10 20 . 30 40 50
TIME FROM DPS IGNITION (SEC)
Figure 4-6. Second DPS Burn Thrust Profile
38
o .-
If 300
170
OZ '-: ; •>=•' " 160
120
80
180 190 230,200 . 210 220
DESIRED APOGEE FOLLOWING SECOND DPS BURN ( N Ml)
250
NOTE:1 . DESIRED STUDY PERIGEE IS 175 N Ml2. ASSUMED IMU DRIFT OF .1 .2 DEC APPLIED
IN DIRECTION OF MAXIMUM ORBIT SENSITIVITY.
Figure 4-7. Effect of Assumed IMU Drift on Orbit Dimensionsfollowing Second DPS Burn
FITH STAGING OCCURS4 HRS 47 MIN 56 SECFROM LIFTOFF
. 200
t '60LU
u
1/1 120
LU
Lu 80a:
• 40
0
i i
i T
r
*TIV
E V
ELO
CIT
Y (
FT
/SE
C)
. '
''
0
A.
' O
. 00
3
O
O
O
_jLUa:
. - ol
- • FIT
I—— PLt
f^- 0
\ STAG IN
ME IMPIN
VELOCITY
X'
/
5 /APS IG
CEMENT E
' DISTAN
^ - - - - - -
^PS SHUTD
NITION-
FFECTS ,
^:E
OWN-*-
/
f
^
2 ' 3 4 5
TIME FROM FITH STAGING (SEC)
Figure 4-8. FITH Staging; Relative Velocity and Distance HistoryDuring first APS Burn"
39
AN
CE (F
T)
7000
IV
LAT
IVE
VE
LOC
ITY
(F
FITH STAGING OCCURS4 MRS 47 MIN 56 SECFROM LIFTOFF
20 40 60 80
TIME FROM FITH STAGING (SEC)
100
Figure 4-9. FITH Staging; Relative Velocity and DistanceHistory to 100 Seconds
170 10 190 200 2JO 220 230 240DESIRED APOGEE FOLLOWING SECOND APS BURN ( N Ml)
2 0
NOTE:DESIRED STUDY PERIGEE IS 175 N Ml
2. ASSUMED IMU DRIFT OF 1.7 DEGREES APPLIEDIN DIRECTION OF MAXIMUM ORBIT SENSITIVITY
Figure 4-10. Effect of Assumed IMU Drift on Orbit DimensionsFollowing Second APS Burn
40
5. NOMINAL"TRAJECTORY DATA
, This section contains pertinent trajectory parameter histories of the
nominal mission profile. These data, in tabular and graphical form, are
based upon the trajectory data presented in Volume II'of this report.
Volume II also contains definitions of the trajectory parameters presented
i n this section. • ' • ' . , -
5. 1 MISSION PROFILE DATA ; :
The time sequence of events for Apollo Mission AS-206A is presented
in Table 5-1. Figures 5-1 through 5-4 present the earth ground track for
the entire mission. Orbital characteristics of the spacecraft coast phases
are presented in Table 5-2. Earth shadow data (daylight-darkness) are
illustrated in Figures 5-1 through 5-4 and tabulated in Table 5-3.
5.2 .TRAJECTORY PHASE DATA
Discrete events summaries and pertinent time history ;data of the
spacecraft position, motion, • and attitude are presented for the major ' '
mission phases as follows:
Mission Phase Table • • Figures
Ascent-to-Orbit 5-4 5-5 through 5-8
Spacecraft Separation 5-5 5-9 through 5-11
First DPS Burn.. ' . 5-6 5-12 through 5-15 :
Second D P S Burn/FITH'Staging/ . . . . . . .First APS Burn . 5.-7 5-16 through 5-19
Second APS Burn 5-8 . . . . 5-20 .through 5-23
Third APS Burn 5-9 5-24 through 5-26
Fourth APS Burn 5-10 . 5-27 through 5-29
The spacecraft attitude angles presented in the figures are refer-
enced to a launch-centered inertial-coordinate system. This-coordinate
system and the spacecraft axis system are illustrated in .Figures 3-7 and
3-8.
41
Table 5-1. Time Sequence of Events
• ' . - - . - . - . . - . - . - Time From LiftoffEvent (hr:min:sec)
Ascenti-to-Orbit
-- Liftoff - - . -. . 00:00:00.0S-IB Inboard Engines Shutdown 00:02:17.6S-IB Outboard Engines Shutdown 00:02:23. 6S-IB Jettison/S-IVB Ignition 00:02:29. 1Jettison Boilerplate CSM 00:02:49. 1S-IVB Shutdown into Elliptical Earth Orbit 00:10:03. 4
S-IVB/Spacecraft Orbital Coast
S-IVB Shutdown into Elliptical Earth Orbit 00:10:03. 4SLA Petal Deployment 00:40:03. 4
Spacecraft Separation
RCS Ignition 00:52:02. 8Sever Restraining Straps 00:52:04.8RCS Shutdown 00:52:20. 8
Spacecraft Orbital Cold-Soak
RCS Shutdown 00:52:20. 8Begin Cold-Soak Orientation Maneuver ' 00:52:30. 8End Cold-Soak Orientation Maneuver 00:53:26.1
First DPS Burn
Maneuver to Pre-ignition Attitude 03:55:32.0LEM Landing Gear Deployment 03:56:32. 0RCS Ignition 03:57:32.0First DPS Ignition 03:57:38.0First DPS Shutdown 03:58:16. 7
Second DPS Burn
Begin Pre-ignition Orientation Maneuver 04:33:13. 2RCS Ullage Maneuver 04:35:13.2Second DPS Ignition 04:35:19.2Second DPS Shutdown/FITH Staging/First APS
Ignition 04:47:56.2First APS Shutdown 04:48:01.5
42
Table 5-1. Time Sequence ,of-.Events (Continued)
' • • - - . . . Time From LiftoffEvent (hr:min:sec)
Second APS Burn ' '
Begin Pre-ignition Orientation Maneuver 06:09:04. 9RCS Ullage Maneuver - : • 06:11:04. 9Second APS Ignition V.. . - 06:11:07'. 4Second APS Shutdown- ; "" ' . 06:17:42.1
Third APS Burn . • ' • : • ' % : • • - • • . . • • : - . . • :
Begin Pre-ignition Orientation Maneuver ' 07:1.0:46. 1RCS Ullage Maneuver " . 07:12:46. 1Third APS Ignition . '" • '07:12:48.'6Third APS Shutdown "• 07:12:53.9
Spacecraft Orbital Cold-Soak • • • • . ' ' . - . - - • • . .
Third APS Shutdown . . 07:12:53.9Begin Cold-Soak Orientation Maneuver : •• 07:13:03.9End Cold-Soak Orientation Maneuver . 07:13:57.6 ,
Fourth APS Burn • ' • ' " ' • '
Begin Pre-ignition Attitude Maneuver 10:58:53.5.RCS Ullage Maneuver ' ' 11:00:53.5Fourth APS Ignition . - • • - . . . - \ • " i i:QO:56'. 6Fourth APS Shutdown -' 11:01:04.3
Final Orbital Coast
Fourth APS Shutdown • ' • - • • • 11:01:04. 3End of Mission Profile 14:00:00.0
•43
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Table 5-3. Spacecraft Earth Shadow Data
Entrance IntoEarth's Shadow
(Time from Liftoff)(hr :min)
00:47
02:16
03:44 •
05:14
06:43 •
08:14 .
09:45
11:17
12:47
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(Time from Liftoff)(hr :min) '
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04:20
05:49
- 07:18 ;
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(min)
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Figure 5-10. Spacecraft Separation; Inertial Velocity, InertialFlight Path Angle, and Inertial Azimuth
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6. MSFN COVERAGE DATA
Spacecraft visibility periods for the MSFN stations presented in
Table 3-1 are listed in Table 6-1. Also presented in this table are the
approximate minimum and maximum slant ranges between the spacecraft
and the MSFN station during each visibility period. Table 6-2 presents
the time intervals in which the spacecraft is not visible from a MSFN
station.. The spacecraft is considered visible when the line-of-sight be-
tween the spacecraft and the MSFN station is 85 degrees or less with
respect to the station vertical and the slant range distance is less than
740 nautical miles. The MSFN coverage during major mission events is
illustrated in Figure 6-1.
Volume III presents detailed tracking time history data for the
ground stations available for support of this mission. These data consist
of range, range rate, azimuth angle, azimuth angle rate, elevation angle,
elevation angle rate, and four spacecraft-to-MSFN station look-angles.
Volume III data are presented as a function of time for each of the" ground
stations and annotated for significant mission events.
67
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73
Table 6-2. Communication Void Intervals
Void BeginsTime from Liftoff
(hr:min:sec)
OQ:14:08
00:22:38 . . . ' . . .
00:57:45
01:47:08
02:31:39
02:56:04
03:18:19
03:33:14
03:45:26
04:05:06
04:30:34
04:54:13
05:08:25
05:23:59
05:40:58
06:06:27
06:30:17
06:44:38
06:59:27
07:16:57
07:42:51
08:07:31
08:21:17
08:33:11
08:50:27
09:05:16
09:18:43
09:33:51
10:13:20
10:41:13
10:55:25
Void EndsTime from Liftoff
(hr:min:sec)
00:18:09
.00:48:52
01:28:45
02:22:24
02:52:23
03:00:21
03:29:18
03:40:02
03:55:31
04:24:32
04:33:13
05:03:55
05:15:02
05:30:38
06:00:14
06:09:05
06:39:43
06:50:51
07:06:36
07:36:44
07:45:40
08:15:43
08:27:18
08:43:05
08:59:27
09:14:19
09:22:05
10:04:00
10:36:13
10:51:38
10:58:54
VoidDuration(min: sec)
04:01
„. ... ._ -26-16 -
31:00
35:16
20:44
04:17
10:59
06:48
10:05
19:26
03:39
09:42
06:37
06:39
19:16 .
02:38
09:26
06:13
07:09
19:47
02:49
08:12
06:01
09:54
09:00
09:03
03:22
30:09
22:53
10:25
03:29
74
Table 6-2. Communication Void Intervals (Continued)
Void BeginsTime from Liftoff
(hr:min:sec)
11:06:12
11:50:50
12:32:15
13:20:49
' 13:25:18
Void Ends .Time from Liftoff
(hr:min:sec) .
11:40:27 ' • ,
,12;26:35 . " ..
..13:15:56
13:21:12
14:00:00*
Void •Duration
: (min:sec)
.. : 34:15
35:45
43:41
00:23
34:42
* End of mission profile.
75
ASCENT-TO-ORBIT
-LIFTOFF
'f1j!!ill
1 j" i
S-IB INBOARD ENGINES SHUTDOWNS-IB OUTBOARD ENGINES SHUTDOWN
• S-IB + ADAPTER JETTISON/S-IVB IGNITION•» JETTISON BOILERPLATE CSM
(-•-INSERTION INTO ELLI
_ dGRAND TURKfl ^
| BERMUDAGRAND BAHAMA |
CAPE KENNEDY f
BERMUDA SHIP 1
|
III I0:00 0:02 0.04 0:06 0:08 0:10
TIME FROM LIFTOFF (HR:MIN)
0:12 0:14
SLA PETAL DEPLOYMENT
II
TANANARIVE!
0:38 0:39 0:40 0:41
TIME FROM LIFTOFF (HR:MIN)
0:42
SPACECRAFT SEPARATION•AHITUDE MANEUVER TOCOLD-SOAK ORIENTATION
KC.1 IU
BEGIN INERTIAL 1ATTITUDE HOLD-.-J
1
MIIIUN
( AUSTRALIA SHIP1
i .'!s SEVER RESTRAINING STRAPSI RCS SHUTDOWN
I I I! 1 i CARNARVON!! 1
1
0:48 0:50 0:52 0:54
TIME FROM LIFTOFF (HR-.MIN)
0:56 0:58
FIRST DPS BURN
-BEGIN PRE-BURN ORIENTATION MANEUVER
iiii1
LANDING GEARDEPLOYMENT ••
j
IRCS IGNITION[K/DPSy ixr i1 1II 1II 1II 1'! I
1 AUSTRALIA SHIP
IGNITION)PS SHUTDOWN
I CARNARVON ||
3:54 3:56 3:58 4:00 4:02
TIME FROM LIFTOFF (HR:MIN)
4:04 4K)6
Figure 6-1. MSFN Coverage Summary
76
SECOND DPS BURN/FITH STAGING/FIRST APS BURN
BEGIN PRE-BURN ORIENTATION MANEUVER
/
f^RCS IGNITION f
I l"
P f TEXA
| | WHITE SANDS
J | GUAYMAS |K GOLDSTONE f
I PT.ARGUELLO ]
CALIFORNIA SHIP|
1
!
f BERMUDA SHIP J
f ANTIGUA |BERMUDA |
I GRAND TURK |GRAND BAHAMA J
CAPE KENNEDY J
EGLIN TJ
S J f .APS SHUTDOWN
J \SI FITH STAGING/r APS IGNITION
1
!4:40 4:45
TIME FROM LIFTOFF (HR*1IN)
SECOND APS BURN
XBEG IN PREBURN ORIENTATION MANEUVER
.—RCS IGNITION
Ih——.APS IGNITION
-APS SHUTDOWN
9 r? '
"
1
[ EGLIN :
WHITE SANDS
GUAYMAS1 h GOLDSTONE
1 PT ARGUELLOB CALIFORNIA SHIP 1
I
TEXAS :
11
6:12 6:14 6:16TIME FROM LIFTOFF (HR:MIN)
THIRD APS BURN
L
7:06
j1t
J
1
lj~~
AUSTRALIA SHIP i
7:08 7:10
;— RCS IGNITIONij— APS IGNITIONft-- APS SHUTDOWN
X !T r\^, _ a.T|lTUI>f MANF^VfR JO|J! ! I COLD-SOAK ORIENTATION
*!' CARNARVON \
:!:i i7:12 7:14 7:16 7:18
TIME FROM LIFTOFF (HR:MIN)
FOURTH APS BURN
BEGIN PRE-BURN ORIENTATION MANEUVER
^ RCS IGNITION
JJAPS IGNITION
^APS SHUTDOWN
APS
f^A
110:57
1
1 110:58 10:59
"1 1III PT
GOLDSTONE
AGUELLO!' 'CALIFORNIA SHIPin
IIK10
11
1 1 K3i 1 1 ai
1
TIME FROM LIFTOFF (HRJUIN)
Figure 6-1. MSFN Coverage Summary (Continued)
77
7. SUMMARY OF TECHNICAL ACHIEVEMENT
This report contains no innovations or improvements involving new
technology, approaches, methods, or patentable ideas as defined in the
contract's "New Technology and Property Rights in Inventions" clause.
79
REFERENCES
1. "Mission Requirements for Apollo Spacecraft Development MissionAS-206A, " TRW No. 2132-H004-RU-000, 1 December 1965.
2. R. K. Petersburg, "Apollo Mission SA-206A Spacecraft PreliminaryReference Trajectory (U), " TRW No. 3300-H007-RCOOO,1 July 1965. (C)
3. "Apollo AS-206A Trajectory Listing, " furnished MSC by MSFCthrough the Guidance and Performance Subpanel.
4. "Apollo Navigation Working Group, " NASA No. 65-AN-l. 0,5 February 1965.
5. "Operational Support Plan for the Apollo 200 Series Missions, "Prepared by the Flight Control Division/MSC, April 1965.
6. A. Kelemen, "The LEM-1 Mission Capability Report NASA MissionAS-206A (U), " No. LED-540-41, 1 November 1965. (C)
7. Minutes of Flight Operations Plan, Meetings 1 through 12.
8. "Apollo Mission Data Specification C Apollo-Saturn 206A (U), "TRW No. 2131-6002-TCOOO, 5 November 1965. (C)
9. "Launch Vehicle Reference Trajectory, " MSFC, to be published.
10. "Lifetime of Near Earth Satellites in Circular or Elliptical Orbits, "Memorandum for Record, OFO (JCB:jec), 1 3 September 1 963. (C)
11. "Apollo Guidance and Navigation, G and N System Operations PlanMission AS-206, " MIT Instrumentation Laboratory No. R-527,November 1965. (C)
81