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TRANSCRIPT
I 3300-HOOT-RC-OOO
Vol. I
1 " _ "
C_D[- 7_7J
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APOLLO MISSION SA-206A
SPACECRAFT PRE LIAIIN._LRY
REFERENCE TRAJECTORY (U)
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ti JULY 1965 4
, Volume I
TRAJECTORY DESCRIPTION
"_ Prepared for
_7 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
MANNED SPACECRAFT CENTER
Contract No NAS 9-2938
Phase II (Apollo) "
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TRWsvSTEMS
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" APOLLO MISSION SA-206A
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.... REFERENCE TRAJECTORY¢
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TRAJECTORY DESCRIPTION
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-" . ..... " - Prepared for
NATIONAL AERONAUTICS ANE) SPACE ADMINISTRATION
- .-_. ,. MANNED SPACECRAFT CENTERContract No. NAS 9-2938
" :: ."". " Phase II (Apollo)
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330_H007-RCO00
Total Pases: ]52
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APOLLO MISSION SA-206A
!SPACECRAFTPRE-LI_fINARY
REFERENCE TRAJECTORY_
It, julY 196s
Volume I
TRAJECTORY DESCRIPTION
Prepared for
INIATIONAL AERONAUTICS AND SPACE ADMINISTRATION
MANNED SPACECRAFT CENTER
Contract No. NAS 9-2938
Phase Ii (Apo!'o)
1i Carl R. Huss
Chief
Flight Analysis Branch
Approved by.___=_i._.
,!1 -__ .... ,^_,0_,;i_1_,Mayer, ....
" Mission Planning and
I Analys|s Division-
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Approvedby . V. Stab,efoj_
ManagerManned Spaceflight
Department
Approved byE. A. Ward
........ Ma nagerMission Planning and
Operations, MTCP
_J SYSTE;'_S
" _ 3300-H008-RC000
Page ii
FOREWORD
This report,'which de/ines the Spacecraft Preliminarym
Reference Trajectory for Apollo Mission SA-Z06A,is sub-
mittedby.TRW Systems to the NASA Manned Spacecraft Center
in partial response to Task A-21 (Establishment of t h e
Reference Trajectory ior Apollo _V_ission SA-Z06A) of the
Apollo Mission Trajectory Control Program (Contract No,
.NASg-293B, Phase II). This report is presented in three
volumes. Volume I sunun_rizes the mission objectives, the
mission guidelines, and the L:put d_ta for the mission simu-
lation and provides a detailed descriptionofthe mission pro-
file. Graphical and tabular :irne history data of spacecraft
attitude, position, motion, and other pertinent trajectory
data are also presented in Volume I. Volume II contains the
trajectory listing of the mission profile, along with the
trajectory print key. Detailed tracking time history data
are presented graphically h _. Volume III.and annotated fo r
significant events.
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CONTENTS
3300-H007-RC000
P_ge iii
LNTRODUCTION AND SUM_V_ARY .................. 1
I. I Purpose. , . . . .......... ....... ......... I
I. Z Scope ........ .............. .......... I
I. 3 Mission Profile Summary .................. I
SPACECRAFT MISSION REQUIRE.'vtENTS ........... 5
2. I Spacecraft Test Objectives ................. 5
Z. I. I Primary .......................... 5
Z.l.Z Secondary ......................... 6
Z. Z Mission Profile Guidelines ................ 6
2. Z. I Launch Vehicle Systems .............. 6
Z.Z.Z Spacecraft Systems .................. 6
SU.V_JkRY OF INPUT DATA .................... 9
93.1 Saturn IB Launch Vch:cle ..................
3. Z Spacecraft (LEM- I) ...................... 16
3.3 .V.SFN Stations .......................... 16
3.4 Earth Constants and Conversion Factors ........ 16
3.4.1 Earth Constar:s .................... 17
3.4. Z Conversion Fac:ors ................. 17
Spacecraft and Reference Coordinate Systems ..... Z33.5
.MISSION ANALYSIS AND DESCRIPTION ............ Z5
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4.
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4.
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4.
1 Saturn IB Ascent to Orbit .................. Z5
Z S-I_,'B/SLA/LEIV[ Orbital Coast .............. Z6
3 Spacecraft Separate-on .................... Z6
4 Orbital Cold-Soak to First DPS Burn .......... Z7
5 First DPS Burn ......................... 30
6 Orbital Coast to Second DPS Burn ............ 30
T Second DPS Burn ........... , ...... . . . . .. 30
8 Orbital Coast to FITH Abort Test ............. 33
9 FITH Abort Test ........................ 33
I0 Orbital Coast to Second APS Burn ............ 34
II Second APS Burn ............. - ....... . • • 39
IZ Orbital Cold-Soak to Third APS Burn . ......... 39
• 4013 Third APS Burn ................... .....
4014 Final Orbital Co st ......................
15 Orbital Lifetime Estimates .................. 40
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3300-H007-RC000
Page iv
NO.*_INAL TRAJECTORY DATA ......... , ..... , . , 43
5. 1 Mission Profile Data ................. . . , .
5. Z Trajectory Phase Data ................. . . . .
6, TRACKL_G AND COMMUNICATIONS DATA .... .......
7. SU3MIfA/KY OF TECHNICAL ACHIEVE.MENT .........
APPENDIX
OPEN-LOOP ATTITUDE A£AA_EUVEK LOGIC .........
43
43
t28
137
,ss
REFERENCES ............................... , , . 140
ABBREVIATIONS . . . eee..e,e..e.ee.eeeeeeeeeeee'e
142
Total Pages: 152
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3300-H007-RC000
Page v
,.i [LLUSTRATIONS- Pag___e
, 1-1 Mission Summary .................... , . , , • 33- I S-IB Thrust Profile . ....................... £2
i 3-2 . S-IB Propellant Weight Flow R_te Profile ......... t 33- 3 Saturn IB Launch Vehicle ........ , ........... 14
• 3-4 Saturn IB Zero-Lift Drag Coefficient (power-on
i and power-of0 , .................. ....... i 53-5 "LEM-DPS Specific Impulse and Propellant %_ eight
Flow Rate ............................... 20
i 3-6 Spacecraft (LEM) " 21
3-7 Spacecraft and Reference System Coordinates ...... 24
6 4-ta S-IVB/LEMvelocityRelative Velocity and.................Distance
to ZZ Seconds ............................ Z8
ll 4-1b S-IVB/LEM Relative Velocity and Distanceto 240 Seconds ....... ' .................... 29
4-2 LEM-DPS Thrust Profiles ................... 3i
4-3 LEM-DPS Propellant Weight Flow Rate Profiles . - , . . 32
4-4a Relative Velocity and Separation Distance
• FoLlowing LEM Staging to APS Shut/own .......... 35
I 4-4b Relative and Separation DistanceFollo_ng Lz_M Staging to 8 Seconds ............. 36
• 4-5a Rela_ve Position Coordinates Following
LEM Staging to APS Shutdown 374-5b Relative Position Coordinates Following
l _ Staging to 8 Seconds .................... 38
5-I Earth Ground Track/Entire Mission Profile ........ 47
5-Z Earth Ground Track/Second DPS Burn ............ 50
5-3 Earth Ground Track/FITH Abort Test ............ 50
5-4 Saturn IB Ascent to Orbit/Altitude, Latitude,
and Longitude ............................ 56-5-5 Saturn ]]5 Ascent to Orbit/Inertial Velocity, Flight
I Path Angle, and Azimuth .................... 575-6 Saturn IB Ascent to 0rbit/Relative Velocity, Flight
• Path Angle, and Azimuth ..................... 58
5 8 Saturn IB Ascent to Orbit/Altitude, Mach Number,
Drag and Dynamic Pressure ................... 60
Saturn IB Ascent to orbit/Pitch Rate, and Pitch
9 Angle of Attack ........................... 615 I0 Saturn IB Ascent to Orbit/Vehlcle Attitude
_Laumch S_te Inertlal) .......................63
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5-15
5-t6
5-17
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5-21
5-22
5-23
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5-25
5-26
5-27
5-28
5-29
5-30
5-3t
ILLUSTRATIONS (Continued)
3300-H007-RC000
Page vi
Saturn IB Ascent to Orbit/Vehicle Attitude(Earth Referenced Rotating) ................... 64
S-IVBISLA/LEM Orbital Coast/A/titude, Latitude,and LongL__de ............................ 66
S-IVBISLAILEM Orbital Coast/Inertial Velocity',Flight Path Angle and Azimuth .................. 67
S-IVB/SLA/L_M Orbital Coast/Vehicle Attitude
(Launch Si_e _ertial) ............ ........... 68
"S-IVB/SI_A/LE__ Orbital Coast/Vehicle Attitude
(Earth Referenced Rotating) ................... 69
Spacecraft Separation/Altitude, Latitude, and
Longitude ............................... 71
Spacecraft Separation/Inertial Velocity, Flight
Path Angle, and Azimuth ..................... 7Z
SpacecraSt Separation�Spacecraft Attitude
(Launch Site inertial) ....................... 73
SvacecraSt SeTaration/Spacecraft Attitude
{Earth Referenced P_otating)................... 74
Spacecraft Separation/Total Acceleration ......... 75
Orbit_ Cold-Soak to First DPS Burn/Altitude,
Latitude, and Longitude ............... . ..... 77
Orbital Cold-Soak to First DPS Burn/Inertial
Velocity, F_ght Path Angle, and Azimuth ......... 78
Orbital Cold-Soak to First DPS Burn/Spacecraft
Attitude (La'_ch Site Inertial).................. 79
Orbital Cold-Soak to First DPS Burn/Spacecraft
Attitude (Ear:h Referenced Rotating) ............. 80
First DPS Burn/Altitude, Latitude, and
Longitude ............................... 82
First DPS Burn/Inertial Velocity, Flight Path
Angle, and A_Ln_.u_.h......................... 83.#
First DPS Burn/Spacecraft Attitude(Launch Site L--ertial)........................ 84
First DPS Burn/Spacecraft Attitude (Earth
Referenced Rotating) ........................ 85
First DPS Burn/Total Acceleration .............. 86
Orbital Coast to Second DPS Burn/Altltude,
Latitude, and Longitude ...... ................ 88
Orbital Co_st to Second DPS Burn/Inertial Velocity,
Flight Path Angle, and Azimuth ................ 89
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ILLUSTRATIONS (Continued)
3300-HOO7-RC000
Page vii
Second DPS Burn/Altitude, Latitude, and
Longitude ................................ 91
Second DPS Burn/Inertial Velocity, Flight Path
Angle, and Azimuth .......... , .............. 9Z
Second DPS Burn/Spacecraft Attitude (Launch
Site Iner tial) ........................... , .... 93
Second DPS Burn/Spacecraft Attitude (Earth
Referenced Rotating) 94
Second DPS Burn/Total Acceleration .............. 95
Orbits/Coast to FITH Abort Test/Altitude,
Latitude, and Longitude ...................... 97
Orbital Coast to FITH Abort Test/Inertial
Velocity, Flight Path Angle, and Azimuth ......... . . 98
FITH Abort Test/Altitude, Latitude, and
Longitude ................................ 100
FITH Abort Test/Inertial Velocity, Flight Path
Angle, and Azimuth ......................... 101
FITH Abort Test/Spacecraft Attitude (Launch
Site Inertial) .............................. 102
FITH Abort Test/Spacecraft Attitude (_arLh
Referenced Rotating) ........................ 103
FIlq-I_bort Test/Total Acceleration .............. 104
Orbital Coast to Second APS Burn/Altitude,
•Latitude and Longitude I06
Orbital Coast to Second •APS Burn/Inertlal
Velocity, Flight Path Angle, and AzL_c.u:h .......... 107
Second A-PS Burn/Altitude, Latitude, and Longitude .... 109
Second AIDS Burn/Inertial Velocity, Flight Path
Angle, and Azimuth ......................... 110
Second A-PS Burn/Spacecraft Attitude (Launch
Site Inertial) .............................. 111
Second _S Burn/Spacecraft Attitude (Earth
Referenced l_otating)......................... 112
Second APS Burn/Total Acceleration .............. 113
Orbital Cold-Soak to Third APS Burn/Altitude,
Latitude, and Longitude ...................... 115
Orbital Cold-Soak to Third APS Burn/Inertial
Velocity, Flight Path Angle, and Azimuth .......... If6
Orbital Cold-Soak to Third APS BurnlS_cecraft
Attitude (Launch Site Inertial) 117• t • • o • o • • • • • • •- • • • •
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ILLUSTRATIONS (Continued)
3300.H007-RCO00
Page viii
Page
Orbital Cold-Soak to Third APS Burn/Spacecraft
Attitude (Earth Referenced l_otating) .............. 118
Third _ Burn/Altitude, Latitude and Longitude ..... 120
Third APS Burn/Inertial Velocity, Flight Path
Angle, and Azimuth ......... , ............... 121
Third A_PS Burn/Spacecraft Attitude (LaunchInertia/) 122Site ........ - ..........
Third _ Burn/Spacecraft Attitude (Earth
.Referenced Rotating) ........................ IZ3
Third APS Burn/Total Acceleration .............. 124
Final OrbitaI Coast/Altitude, Latitude, and Longitude . . 126
Final Orbital Coast/Inertial Velocity', Flight Path
Angle, and Azimuth .................. 127
MSFZNTracking Summary ..................... 135
• Euler Angle Transformation .................. 139
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3300-H007-I_C000
Page ix
• TABLES
Pag_.__ee
Saturn IB Weight Statement ................... 10
Saturn IB Propuls ion Data .................... 11
Saturn IB Event Timing Criter_ ................ 11
LEM-! Weight Statement .................... . 18
Criteria for LEM-RCS Propellant Expenditures ...... 19
Radar Tracking Station Sites and Eq,,_iprnent ........ 22
Ba]/istlc Coefficients ....................... 41
"Orbita/Lifetime Estimates ................... 41
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5-4
5-5
5-6
5-7
5-8
5-9
5-i4
5-i5
5-t6
°5-t7
5-18
5-t9
6--t
6-2
Time Sequence of Events ..................... 45
Orbital Characteristics of the Spacecraft Coast Phases. 51
Earth Shadow Data ......................... 5Z
Spacecraft Body Attitude l_ate History .... "........ 53
L_'vI-P, CS PropelL_nt E_v_end-bares .............. 54
Saturn T_ Ascent to Orbit/Discrete Events Summary... 55
S-IVB/SLA/LEM Orbital Co:,_/Diserete Events
Summary .......................... •..... 65
Spacecraft Separation/Discrete _vents Summary ..... 7_
Orbital Cold-Soak to First DPS Burn/Discrete Events
Sunu_ry ............................... 76
First DPS Burn/Discrete Events Sun-unary.. ....... 81
Orbital Coast to Second DPS _urn/Discrete Events
Summary ............................... 87
Second DPS Burn/Discrete Events Summary ........ 90
Orbital Coast to FITH Abor: Test/Discrete Events
Summary ............................... 96
FITH Abort Test/Discrete Events Summary ........ 99
Orbital Coast to Second A-_S 3urn/Discrete Events
Sun_ry ............................... 105
Second AIDS Burn/Discrete Y_ven_ Summary .......... 3
Orbital Cold-Soak to T]_rd .%PS 3urn/Discrete Events
Summary " 114• • . , . . • • . • . • • • • * • • • • * • • * • * * • • •
Third APS Burn/Discrete Events Su2nmary. ........ 119
Final Orbital Coast/Discrete Events Sun,_-n&ry ...... 125
MSFN Tracking Coverage .................... 129
Communications Void Intez'vals 133
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I. INTRODUCTION AND SUAtMAKY
i.i PURPOSE
3300-H007-KC000
P_ge i
The Spacecraft Preliminary Reference Trajectory defined in this
document is designed for the unmanned Apollo Mission SA-Za6A. It is a
combined launch vehicle and spacecraft trajectory prof_e "_th_tis intended
to satisfy the mission's primary spacecraft objectives (-_[eference i)with-
out violating any of the launch vehicle and spacecraft cons_ra_s or the
mission guidelines. Other than the removal of the lonE-duration cold-soak
requirements, the basic trajectory profile is similar to that presented in
the Prelimirm.ry Mission Profile, Reference 4. The purpose of this report
is to improve upon and expand the scope of the trajectory proFde presented
in Reference 4 while complying with Reference I.
i. Z SCOPE
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This report is presented in three volumes. Vol,_e ! subz_..its the
spacecraft mission requirements, summarizes the inpu*, da:a used in the
mission simulation, describes the major phases of *-he trajectory, and
gives the trajectory analysis for applicable phases. It contains graphical
and tabular time history data of the spacecraft attitude, position, and
motion. Spacecraft separation characteristics and tracking station visi-
bility data are also presented in this volume.
Volume II of this report contains the trajectory _s_Jing of the
miss ion simulation.
Volume III presents detailed tracking time h/story data for the ground
stations available for operation on this mission. These data consist of
range, range rate, azimuth angle, elevation an_le, and t_vo spacecraft-to-
MSFN statio n look angles and are presented as a function of tL-ne for each
of the ground stations. The times of significant events are noted on these
plots.
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i. 3 MISSION PROFILE SUAd_MAKY
Apollo Mission SA-Z06A, currently planned for *.he second uuarter of(
1967, will be the first launch of a complete LEM spacecraft. For mission
simulation purposes, the launch on an azimuth of 7Z ° frcunu true North is
assumed to occur at 13:00 GAdT, April i, from Launch Conuplex 37B of
the Kennedy Spaceflight Center.
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Page 2
Major events of the mission are illustrated in Figure I-1. The
mission has been divided into i4 major phases:
1. Saturn IB Ascent to OrbitI
2. S-IVB/SLA/_M Orbital Coast
3. Spacecraft Separation
4. Orbital Cold-Soak to First DPS Burn
5. First DPS Burn
6. Orbital Coast to Second DPS Burn
7. Second DPS Burn
8. Orbital Coast to FITH Abort Test
9. FITH Abort Test*
I0. Orbital Coast to Second APS Burn
it. Second A_PS Burn
. 12. Orbital Cold-Soa:_ to Third APS Burn
i3. Third APS Burn
i4. Final Orbital Coast
The Saturn IB launch 7"_ase includes the burn of the S-IB stage and
the burn of the S-IVB stake. The dummy CSM is jettisoned by fir!no_ the
LES jettison motor at a po_: where the dynamic pressure is less than one
pound per square foot.
S-IVB cutoff occurs at an altitude of 85 nautical miles and a z_.ro degree
flight path angle, with the ve!0ci.'7 necessary for an elliptical orbit insertion,
having an apogee altitude of !20 nauticaI miles.
The spacecraft is separated from the S-IVB/SLA combination on
the first orbit by the LE_'---_CS thrusters while in sight of the Carn_rvon
tracking station.
The first DPS burn is performed on the third orbit after the space-
craft has been subjected to an attitude-hold cold-soak (+ X-a:ds normal to
the ecliptic) for approxLn_-ztely 3 hours.
The second DPS burn and the FITH abort test are performed over
the United States at the end of the third and fourth revolutions, respectively.
SThis test includes a third DPS 5urn followed by a FITH abort sin-.ulatlon
(LEM staging/first APS burn).
m m i ml m m m m m m m m m m m-_ .......... !.... I .... j.% .... I
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.:i} _ ;jTWo short duration AI_ burns are then simulated, the first occuring
" " 20 minutes after completion of the FITH abort test and the second one
..... occuring approximately Z. 5 hours after the first. During the Z. _5-hour" . . . . ._
orbital _oast between the short duration A_S burns, the L_-M ascent stage
"+_X-axis is aligned normal to the ecliptic for the second time.
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,.- . " _,nn..o_ RC0005
Z. SPACECRAFT MISSION R.EQUIR.EMENTS
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- The spacecraft test objectives presented here were taken from
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Reference I. --: • .. . _ .... - ,. ..... .- .--
i -- '2. i,i Primary " _. . • : -; :a) Verify LEM subsystems operation after launch vehicle boost
and during and after LEM propulsion system operation.
c)
d)_
Evaluate Flight Control Systems (Guidance and Navigation --
Stabilization and Control -- Reaction Control System) per-
formance and operation at design inertias.
Demonstrate landing gear deployment and determine thermal
distribution resulting from engine plume impingement.
Determine performance and operational characteristics ofthe Electrical Power System (EPS), Environmental Control
System (ECS), and operational instrumentation subsystemsin earth orbit.
e) Determine LEM communications subsystem performance,
operation, and _v_anned Space Flight Net (_MSFN) compatibility.
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g)
Evaluate DPS and APS propulsion subsystems operation
following orbital soaks, including throttle and gimbal control,and demonstrate DPS and A_DS restart.
Demonstrate Fire-ln-The-Hole (FITH) abort and evaluate the
in-flight dyna.-_ics (staging characteristics), pressure distri-
-bution, and Lhermal distribution of the ascent/descent stages
during staging.
h}_ Demonstrate LEIv_ structural integrity, and determine ascent/
descent stage interaction loads, LEN!/SLA interaction loads,
and dymamic loads on pressurant storage and ascent/descent
" stage engine propellant tan.ks.
............. i) ....Evaluate performance and operational characteristics of RCS
in earth orbital environment.o
j) Demonstrate ullage settling time for APS and DPS operation.
k) Determine vibration environment in critical equipment areas,
including engine induced vibration environment during APS
•and DPS operation.
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_ _-. -'_ ...... -_ a) " Demonstrate DPS and APS operation at low.propellant . -
__ quantitie s. - -- .... "-
- / . ..... :....... h) :' Demonstrate operation of the LEM Mission Programmer (LMP).
"-" ".... Z, Z MISSION PROFILE GU!DEL_ES
The following mission profile guidelines for this Preliminary Space-
craft Reference Trajectory kave been compiled from data supplied by NLSC
and from_References i, 2, 4, and 7.
Z. 2. I Launch Vehicle Svste.-z'-_s
" _ _ __at} Launch azimuth of 7Z degrees.
b) The Launch Escape Subsystem (Jettison motor) will be
- utilized to separate the dummy CSI_ from the S-IVB/SLA/LEM
. combination at a point where the dynamic pressure is less
than one pound per square foot.
c) The S-IVB/SLA/LE3.-: combination is to be inserted into an
orbit with conditions similar to Mission 207, but optimized
as to altitude and eccentricity for communications, ground
-control and ground ::'..onitoringaspects of the mission.
|- -d) Guidance comrnand angle rate limitation is to be one degree
• per second in pitch and yaw.
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e) Approximately one orbit of S-IVB stabilization to provide for
LEM subsyste:._s checks, and to provide a stable S-IVB/SLA
" " '" platform from wkich to separate the LEM.
Spacecraft Systems
a) Separation of LE3J from S-IVB/SLA using LEM-RCS thrusters
and deployment of iand-:ng gear during second orbit.
b) LEM attitude mane=-:er rate limitations (in the automatic
mode) is to be !0 de,tees per second in pitch and roll, and 5
degrees per second i."-yaw.
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c)
d)
e)
LEA4 orbital alti.'_:deis not to exceed 300 nautical miles
(communications Hmitation).
The predicted orbital lifetime for the spent descent and ascent
stages is not to exceed three months (also see Reference 19).
The coast times between propulsion tests should be used, as
required, to optin-/ze the ground coverage of the mission.
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Q - Backup ground command of S-I_,'BILEM separation is
- : : k g) The third DPS burnand the first A_PS burn should have good,
- ! -_ ,-T._: - continuous ground coverage.
.... , . .. •
•,..•J:....=_-.__-i/-_h) . The FITH staging demonstration should be positioned so that _
at least three ground stations, with data record capability,• [ can receive these data.
.... i) Orbital soaks are required prior _-o each APS and DPS burn.
_ These soak periods are described below:
Coast for approximately 4 hours with the LEM
X-axis oriented perpendicular to the ecliptic
(when not in the •earth's shadow) prior to firing
the descent stage engine.
Coast, any orientation, for approximately 60
minutes between the first and second descent
stage .engine burns.
Coast, any.qrientation, ZO -+ Z n/mutes between
the first and second ascen_ stage engine burns.
This time interval is cri'.icaI s_nce it is requiredto demonstrate APS restart under maximum heat
soak back conditions.
• Coast for approximately 3 hours with the ascent
stage X-axis oriented perpendicular to the
ecliptic (when not in the ear_-.'s shadow) between
the second and third ascent stage engine burns.
[
A.PS and DPS tests are required as follows:
I
/
• First DPS burn: 25 seconds at t0 percent thrust,
followed by a rapid rise to full thrust and 7 seconds
at I00 percent thrust.
-e--Second DPS burn: 25 seconds at tO percent thrust,
with a rapid ri._e to full ti-.r_st,"-- then a 38S-second
continuous burn with the dnrust decaying li_ear'y
from 100 percent to 90 percent. Decrease thrust
from 90 percent to 45 percent and burn for 115
seconds. Then conduct rando..-r,throttling between
I0 percent and 50 percent _hrust for Z05 seconds.
• Third DPS burn/FITH staging/first APS l_urn:
:_ 25 seconds of DPS firing at I0 percent thrust, a
_ •rapid rise to full thrust, and Z seconds at maximum
thrust. FITH staging, followed by an APS burn
with a duration to ensure propellant depletion after
completion of the third A-P5 burn.
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"_ = _ '_*_*:;_"_o - APS burn" 5 _seconds.
6.
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3. SUALMA_Y OF INPUT DATA
The input data inthis section were extracted from the references and
also include data agreed upon at several technical coordination meetings
between _[SC and TP_W personnel. These data include all quantitative
specifications on the hunch vehicle, spacecraft, and MSFN stations, and
form t_e basis for the Spacecraft Preliminary l_eference Trajectory in
support of Apollo Mission SA-Z06A.
3. I SATURN IB LAUNCH VEHICLE
Data necessary to adequately describe the launch vehicle were
obtained from References 3, 4, 5, 7, 13 and 14 and supplemented by data
from MSC/TRW technical coordination meetings. These launch vehicle data
are included only for completeness and should not be used as official launch
trajectory data or event times.. The official launch vehicle data and launch
trajectory will be published by the MSFC.
A brief weight statement of the Saturn IB launch vehicle is presented
in Table 3-I. The weights are given in a manner essentially equivalent
to their chronological disposition during the mission.
The propulsion characteristics are presented in Table 3-Z. The
operation of the S-IVB stage is divided into three constant-thrust, constant-
flow-rate phases. These phases, listed in order of occurrence, are:
A short duration, nominal thrust, nominal specificimpulse phase.
A high thrust, low specific impulse phase.
A low thrust, high specific impulse phase.
The launch vehicle event timing criteria used in the trajectory gen-
eration are presented in Table 3-3.
The Saturn IB launch vehicle is illustrated in Figure 3-3, and the
"zero-lift drag coefficient (power-on and power-off) data are presented in
Figure 3-4. An aerodynami c reference area of 360.24 square feet was
used.
static atmosphere models are used in the ascent trajectory
simulation. Below an altitude of 35 km (I14,830 feet), the Patrick AFB
"atmosphere (Reference 10) is used, and between 35 km and 400,000 feet,
the U. S. Standard Atmosphere of i96Z (Reference il) is used. No attempt
has been made to remove the small discontinuity between the two models
at 35 kin.
.
Table 3-I.
Event
S-IB Ignition/Liftoff
Saturn IB Weight Statement
Losses .(,lb)
.- -. .. S-IB Impulse Propellant
I S-IB Inboard Engines Cutoff
86t_829
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Event
Weights (Ibl
-1,297=088
435, Z59
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S-IB Outboard Engines Impulse andThrust Decay Propellants
S-IB Outboard Engines Cutoff
Spent S-IB"
S-IB/S-IVB Interstage Adapter
S-IVB Engine Ignition
S-IVB Impulse Propellant
Thermolag and Ullage Cases
24,515
98,826
7,000
227,824
235
4i0,744
304,918
Durruny CSM/LES 9,540
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S-IVB Engine Cutoff
Spent S-IVB
Consumable Propellants l_emaining*
Instrument Unit
Spacecraft LEM Adapter
25,535
t,494
4. i50
3,600
67, 3t9
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Spacecraft (LEM-1) in Orbit
* Includes flight performance reserves.
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Table 3-2.
S-IB Stage ........
(See Figures 3-1 and 3-Z for thrust and propellant weight flow
rate profiles, respectively).
Saturn IB Propulsion Data,, m
_ - . . _- ...: -. ......
5.0 5.5 4.7
10.00 285.33 t52.78
Z05,000 230,000 t90,000
•48t. ZZt 543. 607 444. 476
426. 0 423. I 427.5
S-IVB Sta._e
ProErammed Mixture Ratio
Duration (sec)*
.Th+-_t (Ib)
Propellant Weight Flow Rate (lb/sec)
Specific 1_mpulse (sec)
* Total burn duration of the S-IVB stage is 448. I i seconds.
Table 3-3. Saturn IB Event Timin_ Criteria
!
!
Event
Sa._rn IB Liftoff
Pitch Over/Begin Gravity Turn
End Gravity Turn
S-_B Inboard Engines Cutoff
S-]30u*.board Engines Cutoff
S-_ Ignition
Thermolag and Ullage Cases Jettison
Dununy CSM/LES Jettison*
S-IVB Cutoff
Timing Criteria
t0
t + t0 secondso
t I - Z seconds
t t
tZ (t I + 6 seconds)
t3 (tz + 5.5 seconds)
t3 + iO seconds
t 3 + 10 seconds
t 4
I ...
Dynam/c pressure equal to approximately O. 98 Ib/ft z.
r -,f
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1.8
1.6+
i.4
1.2
1.0
0.8
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0
+
INBOARD ENGINES SHUTDOWN/BEGIN DECAYI
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OUTBOARD ENGINES SHUTDOWN/BEGIN DECAY •l I
h
I,
0:20 0:40 I:00 1:20
0.2
.,er"-11
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.<X
.... O:00
.... < ..
- .,,. ,7./.,7 +
i I
1:,¢0 + 2:00 2:20 2:,_g
TIME FROM LIFTOFF (MIN:SEC)
Figur e 3-1., S-m Thrust Profile
_.. .+
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1 tINBOARD ENGINES SHUTDOWN/BEGIN DECAY
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OUTBOARD ENGINES SHUTDOWN/BEGIN D':-CAY
1tI!
i ,,,
O:00 0.20 0.40 1:00 1:20 l:40
TIME FROM LIFTOFF (MIN:SEC)
1
2:CO 2:20 2:40
//
FiEure 3-Z. S-IB Propellant Weight Flow Rate Profile
__ _ -sso
Page 14
/A_ 15o00,
•n. " t " -
I : 349.421
, II 295.723
I HORIZON
336:000 SENSOR SLA
• .STA 1780.059 /
f HELIUM TUNNEL--.,,.._| STA 1662.859
I _.oooI SYSTEMS TUNNEL_ _ 6! !
/ AUXlUARY TS'':": " S-.iVB
ULLAGE ROCKET GIMBAL STA I
' j" IB
11O0.000
_ . ri_ STA - 1.000 "--_
I NOTES:
I.2.
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srA s.....,,,
_&__,TA2_o.5._ }STA 2189.859 "----------1.33°00'154 _. _i (STA 2034.859
DUMMY CSM
ALL VEHICLE STATIONS AND DIMENSIONS ARE IN INCHES
SATURN REF: MSFC DWG 10M03544,, REV, F
Figure 3-3. Saturn IB Launch Vehicle
- . ..
INSTRUMENTUNIT
i °
- 3300-H007-RC000
e i6
_.2 (LEM-I)
SPACECRAFT
The spacecraft weight breakdown was obtained from Reference 7 and
is based on an inserted payload weight of 36, 140 pounds. This 640-pound
i weight of 35,500 pounds is the expectedincrease from the LEM control
increase in payload capability by ".'-sertion into the nominal elliptical orbit.
i Spacecraft propulsion characteristics and LEM-Reaction Control System(RCS) propellant expenditure criteria were obtained from References 6 and
i 7 and meetings with MSC personnel.
The spacecraft weight statement is presented in Table 3-4. Flight
i performance propellant reserves equal to one percent of the consumablepropellants are assumed. The criteria for establishing the LE.V.-RCS
i propellant expenditures are presented in Table 3-5. The ascent stagepropulsion system is characterized by a vacuum nominal thrust and
propellant weight flow rate of 3,500 pounds and i i. 45 pounds per second,
i respectively, stage propulsion system are presentedThe descent data in
Figure 3-5. An illustration of the spacecraft is shown in Figure 3-6.
i 3. 3 MSFN STATIONS '
The MSFN stations Lhat are p_'anned" to be available for support of this
I mission, their locations, and equipment available were obtained from
Reference 18. These data are surru-narized in Table 3-6. Locations for
three of the five Apollo tracking ships available for support of t_h/s mission
• are also indicated in Table 3-6. One ship has been placed near *..he western
I coast of Australia,' one off the wes:ern coast of the continental United States,and the third near the western coast of A/rica.
I The station coordinates given are based on a Fischer ellipsoid. This
model is described by: ...............
a = equatorial earth radius = 6378166.000 meters (exact)" b = polar earth radius = 6356784.284 meters
f = flattening = 1/298.30
• The altitude is referenced to the ellipsoid and includes a geoidal separation.
3.4 EARTH CONSTANTS AND CONVERSION FACTORS
l_l - The following earth constants and conversion factors (Reference i5)
have been used in the generation of the Spacecraft Preliminary Reference
Trajectory.
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3.4. 1 Earth Constants
R.otation_l rate
.°
Equatorial radius
Average radius
Gravitational parameter (_e)
Coefficients of potential harmonic s
J term (second harmonic)
H term (third harmonic)
D term (fourth harmonic)
3300-H007-RC000
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o
O.
O.
2. 092573819 x lO 7 it
2. 090984t x 107 it
37526902 x iO "3 rad/min _
4t7807416 x 10 -2.deg/sec
729Zi1504 x 10 -4 rad/sec
5. 53039344 x i0 "3 er3/min 2
11.46782384 x 103 er3/day 2
3.986032 x t05km3/sec2
1.407653916 x 1016 ft3/sec 2
1.62345xi0 "3nd
-0.5T5 x iO"5 nd
0.75T5 x iO"5 nd
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Earth flattening (f)
3.4.2 Conversion Factors
Kilometers per foot
Kilometers per nautical mile
Feet per nautical mile
Weight-to-mass ratio
Mas s-to-weight ratio
Feet per earth equatorial radius
Nautical mile per earth
equatorial radius
i/29S. 30nd
0. 304S x 10 -3 km/ft
t. 85z _m/n mi
6076. 115486 ft/n rni
32. 17404856 Ib/slug
O. 03,,080950 slng/Ib
2. 092573819 x 107 ft/er
3443. 93358 n rrd/er
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LEM-i Weight Statement
Weight (Ib 1
:000
I " " Table 3-4. LE: i
!I LEM-I in Orbit -
Descent Stage
I Inert Weight I 4, 6Z3
Usable DPS Propellants z 17, O50
| _DPS Performanc e Re s erve s t 72
I Ascent Stage
Inert Weight i 5,104
Usable APS Propellants 2 4,965
3A/_SPerformance Reserves 50
|Usable RCS Propellants 4
2t, 845
576
10,695
32,540
II
I
Includes dry weight and trapped fluids.Z
Off-loaded by 133 pounds (fuli-tank capacity is 17,355 pounds).3
Approximately one percent of propellants available.4
From Keference 7.
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_ . Table 3-5. Criteria for LEI_-KCS Propellant Expenditures
i Propellant Expenditure
I RCS operation CriteriaSpacecraft Separation I. 25 Iblsec
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Ullages Preceding
t) DPS Operation
2) APS Operation
Attitude Holds (± 5 deg deadband)
i) During LEM Coast
2) DuringAscent Stage Coast
Attitude Holds (-+ 0. 3 deg deadband)
I) D_ing DPS Burns
2) During APS Burns
3) During FITH S_ging
Attitude .Orientation _aneuver*
l) LEM
Z) Ascent Stage
1. Z5 lb/sec
1.25 lb/sec
i 0.5 Ib/burn
0. i5 Ib/sec of burn
10b
t7.6 lb/rr_neu_rer
4. I lb/m_neuver
I
*Attitude n_aneuver about all three axes.
3300-H007-R.C000
3300-H007-RC000Pa.ge Zt
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RENDEZVOUS
RADARA
U,P_-_ DOCKINGTUNNEL
:KINGWINDOW
ASCENT STAGE
VHF ANTENNA (2)
IiI
S-BAND INFLIGHT
ANTENNA (2_RCS THRUSTER
ASSEMBLY
FORWARDTUNI RCS NOZZLE
III
IIII
I
FROM REFERENCE 12)
DESCENTSTAGE
DESCENT ENG_N_ LANDINGSKIRT G EA.'I.
Figure 3-6. Spacecraft (LEM)
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3. 5 SPACE_T AND REFEI%ENCE COOI%DINATE SYSTEMS
The spacecraft attitude is measured by the pitch, yaw, and roll
angles required to rotate from the reference system to the current space-
craft orientation. The reference coordinate systems are illustrated in
Figure 3-7 and defined below.
Earth Referenced Rotating Coordinate System, XI_-YR-ZK :
Right-handed, orthogonal system centered at the vehicle in which theA
positive D( axis extends downrange in the direction of moron and lies in ther A
the plane of the horizon, the positive Y axis extends upward along ther ^
geocentric radius vector, and the positive Z axis extends to the right in• r
a direction orthogonaI to the downrange direction.
Launch Site Inertial Coordinate System, XI-YI-ZI :
Right-handed, orthogonal system in which the origin coincides with^
the launch site, the positive X i axis extends downrange in "-he direction ofA
the launch azimuth and lies in the plane of the horizon, the positive Yi axis
extends upward along the geocentric radius vector at liftot_¢, and the positiveA
_q axis extends to the right in a direction orthogonal to the launch azimuth•
i•
3300--H007-RC000
LAUNCHAZIMUTH
DIRECTION
OF MOTION
• Spacecr'_ft and Reference System Coordinates
-. -.- • .
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4. MISSION ANALYSIS AND DESCR/PTION
The Spacecra_ Pre!!rninaryReference Trajectory for Apollo Mission
SA-Z06A is designed to meet the test objectives of Section Z.i. The Mission
Profile Guidelines af Section 2. Z are followed except for those officially
changed by References Z and 7. To satisfy these objectives and guidelines
and to determine values of the free variables, a certain amount of trajectory
analysis was performed. The results of this analysis, along with a descrip-
tion of the resulting mission profile, are given in this section.
4. I SATURN IB ASCENT TO ORBIT
Launch of Apollo Mission SA-Z06A will occur from Launch Complex 37B
of the Kennedy SpacePC/ght Center during the second quarter of 1967. The
geodetic coordinates of *..hehunch point are 28.531856 degrees North
latitude and 80. 56495Z degrees West longitude. For the trajectory simula-
tion. launch was ass-_-_.ed to occur at 13:00 hours GMT (08:00 hours EST)
on i April 1967.
The Saturn _ ascen_ to orbit phase is initiated by a 10-second vertical
rise followed by a 0.15ZZ degree kick (an instantaneous rotation of the vehicle
attitude and velocity v__ctor) into a IZS-second gravity turn trajectory with a
7Z-degree azimuth hea;./ng. The inboard engines are shutdown approximately
6 seconds prior to S-_ en_.nes cutoff. Following a 5.5-second coast from
S-IB cutoff, the spe-_t S-_ and the interstage adapter are jettisoned. In the
simulation, S-I'VB e-__-"-in-eig'_on also occurs at this time and a pitch rate
of 0. 9904 degree per second downward is initiated. This high pitch rate
steering is terrr_-_:ed 9.00 seconds after ignition, and a low pitch rate of
0. 0765 degree per second do_nward is initiated. Ten seconds after S-rVB
engine ignition, the du.___y CSM is jettisoned by using the LES jettison motor.
This occurs at a ;:-int _:_ere the dynamic pressure is approximately 0.98
pound per square foe:. Reference 17 states that the dummy CSM/LES is
approximately 73 feet away from the thrusting S-rVB/LEM at the end of
tower jettison motor :hrus _ting. Also, the separation velocity remains
positive and the separa_on distance keeps increasing. The thermolag and
ullage cases are also _ettisoned at this time. The low pitch rate steering
is maintained until S-rV'B eagine cutoff at approximately 10 minutes after
lifto_.
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The value of the kick angle and the magnitudes of the two pitch rates
were determined by iteration techniques so that the following conditions
would exist at S-IVB engine cutoff:
i) Inertial velocity of 25,694. 78feet per second.
Z) Inertial flight path angle of zero degrees.
3) Altitude of approximately 85 nautical miles.
4) S-IVB/SLA/LEM weight at insertion of 67,319
pounds°
Conditio;Is i), Z), and 3) result in S-IVB cutoff at an altitude of 85.6 nautical
miles, a zero degree flight path angle, and the velocity necessary for an
elliptical orbit insertion _u_'than apogee altitude of i i9.4 nautical miles.
*(y 4.The inserted we1_h, from condition 4) is consistent with the launch
vehicle capability as extracted from Keferences 3, 7, and 14. Assuming
a flight performance propel/ant reserve of i, 494 pounds at S-IVB cutoff,
the allowable spacecraft _eight is 32,540 pounds in orbit. Table 3-I
presents a more complete breakdo_-n of the inserted weight.
4. Z S-IVB/SLA/LEM OR_BITAL COAST
Ten seconds after orbital insertion, the S-IVB/SLA/LEA[ combination
is maneuvered at a 0.5 de_ree per second rate until the S-!VB +X-axis
lies in the plane of the local horizontal and the - Z-axis is along due geo-
centric radius vector. This attitude is maintained in order to provide a
stable platform from which to separate the spacecraft. The duration of
this orbital coast (insertlo-_. to spacecraft separation) is 45 minutes and
48.4 seconds.
4. 3 SPACECRAFT SEP_3.ATION
The spacecraft separa_on events consist of the Spacecraft LEM
Adapter (SLA) petal dep!o_._r:-._ent,separation of the spacecraft by firing the
LE2_-RCS +X thrusters for IZ seconds, a coast for 8 seconds, followed by
deployment of the LEA[ landing gear.
• The in-orbit position of this sequence of events is near apogee in the
first orbit, rather Lhan during the second orbit as suggested in Section 2. Z.
This selection allows continuous tracking coverage for the events from the
Carnarvon ground station, which ,has ground command capability. Apollo
tracking ship No. f has been located near the western coast of Australia
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Page Z7
in this profile to provide backup coverage for these events and for the sub- ..
sequent first DPS burn.
Separation characteristics during the first several minutes after
separation are illustrated in Figure 4-1. These data are based upon point
mass simulations.
4.4 ORBITAL COLD-SOAK TO FIRST DPS BURN
Approximately 30 seconds after initiation of the LEM lan :ding gear
deployment the LEM is commanded to perform a 53.6-second n-.a:__uver to
alien the +X-axis (yaw axis) normal to the ecliptic and the + Z-axis (roll
axis) toward the sun. $ The sun's position relative to the earth is dependent
upon both the launch date and tLnae of day. The spacecraft orientation with
respect to the earth, in this case, is also launch time and day dependent.
The general attitude maneuver logic used to simulate the spacecr-=ft atti-
tude orientation is summarized in the Appendix. Following this attitude
maneuver, the spacecraft is put into an attitude hold mode ( -+5 degrees
deadband) and maintains this inertial attitude for approximately Z hours
and 57 minutes. In the simulation, the spacecraft attitude dr_ted approxi-
nlately 0. I degree, with no maneuvers, during this cold-soak (see Figure
5-Z3).
After this orbital coast in the attitude-hold mode, a spacecraft
orientation maneuver to the desired DPS ignition attitude is initiated. This
maneuver takes approximately 4i. 5 seconds. The spacecraft holds this
t v_'eattitude until the completion of the first DPS burn. The total i_v.. duration
from spacecraft separation to the RCS ullage maneuver prece;_ir._ "..hefirst
DPS burn is approximately 3 hours and 5 minutes. The dura:io= of this
coast is slightly less than that suggested in the Mission Profile Guide_2unes.
This selection was made to allow the first DPS burn to occur w'-_;_'ethe
spacecraft is in sight of the Carnarvon tracking station, which has the
necessary ground command capability.
*In the automatic mode, the spacecraft is capable of executing a.'_itude rates
up to I0 degrees per second in pitch and roll, and 5 degrees per second in yaw.
3300-H007=P,.C000
Page 28
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- ,, L _
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_ , I ,,'f 1 iJ / j SPACECRAFT SEPARATION
200,. L, / J OCCURS 55 MIN 48.255 SEC--
/ _" J FROM LIFTOFF
-r ! i i0_ / , ,
. 0:00 . . 0:40 ._.1:20 2:00 2:40 3:20 4:00
TIME FROM SEPARATION (MIN:SEC)
4-lb. S-IVB/LEM Relative Velocit 7 and Distance to 240 Seconds
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4.5 FIRST DPS BURN
The first DPS ignition is preceded by an 8-second RCS ullage ma-
neuver. KCS ignition occurs Z minutes after the C_rnarvon tracking station
acquires the spacecraft on the third orbit. The Apollo ship used to track
the spacecraft separation events is also tracking during this event. The
constant spacecraft inertial attitude during the ullage and the DPS burn
(see Figure 5-Z7} increases the orbit perigee altitude by approximately
17 nautical miles.
The first DPS burn consists of 25 seconds at 10 percent thrust,
followed immediately by 7 seconds at 100 percent thrust. Thrust and pro-
pellant weight _ow rate profiles for this burn, and for Lhe subsequent DPS
burns, are shown in Figures 4-2 and 4-3, respectively.
At DPS shutdown, the spacecraft is on an orbit characterized by
perigee and apogee a!ti_udes of if0.0 and 155.8 nautical miles, respectively.
The resulting orbital period is 89. 3 minutes.
4.6 ORBITAL COAST TO SECOND DPS BURN
The spacecraft coasts in orbit, with no attitude constraints, for
approximately 28 minutes. The spacecraft is being tracked approximately
Z0 minutes during this coast. It is expected that certain KCS tests will
be performed during this coast. Xo attempt has been made to simulate
these various tests in t/Ris profile; however, a certain portion of the RCS
propellants available have been allocated for this phase of the mission
(see Table 5-5). .After this coast, a maneuver is initiated to orient the
spacecraft to a desired inertial attitude (see Figure 5-34). This maneuver
takes approximately 39 seconds. The LEM holds this a_itude to the second
DPS burn ignition. The total time duration be_veen the first DPS burn
shutdown and the ullage n_._aneuver preceding the second DPS burn is
approximately 33 minutes and 20 seconds. This is slightly shorter than
that suggested in the k[ission Profile Guidelines, but it is necessary in
order to achieve the tracking required for the second DPS burn.
4, 7 SECOND DPS BURN
An 8-second I_C$ ullage maneuver precedes the second DPS ignition.
"RCS ignition occurs Z minutes after Point Arguello tracking acquisition,
and within sight of Apollo tracking ship No. Z capable of ground-commanding
i ..... 3300-H007-RC000
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the spacecraft, if necessary. The spacecraft is tracked continuously during
the 730-second burn by l l tracking stations across the continental United
States ar, d down the Eastern Test P_ange (ETK).
Most of the AV available from this burn (approximately 7'000 feet
per second) is d/ssipated out of the orbit plane. This is done b> selecting
an inertial attit_._de at ignition (see Figure 5-34) and a spacecraft roll rate
of 0. 0166 degree per second, which places the LEM on an orbit with a
perigee altitude of 140.9 nautical miles and an apogee altitude of 223.7
nautical miles. The period of this orbit is 9t. i minutes with an inclination
of 3I. 425 degrees. This orbit increases the tracking duration of the ground
stations on subsequent revolutions.
The thrust and propellant weight fIow rate profiles for this burn are
shown in ='i_es 4-2 and 4-3, respectively.
4. 8 ORBITAL COAST TO FITH ABOKT TEST
The LEM coasts in orbit with no attL_ude constraints for approximately
I hour and I8 :ninutes. The spacecraft is being tracked approximatel Y
30 m_uu_.es during this coast. .Tt 1:_ expected that various tests of the I_CS
will be continued during this coast period. No attempt has been made to
sin_ulate these tests in this profiIe; however, a certain portion of the RCS
propellants available have been allocated for this phase of the mission
(see Table 5-5). Alter this coast, a 7-second orientation maneuver is
initiated to achieve the desired inertial attitude for the FITH abort test.
The LE_ holds this attitude for 7 minutes and t2 seconds.
4. 9 FITH ABOKT TEST
An 8-second I_CS ullage maneuver precedes this phase of the mission.
RCS ignition occurs approximately 200 seconds after Point Arguello tracking
acquisition. ApoLlo tracking ship No. 2 will provide the necessary ground
command capability. The FITH abort test consists of a Z7-second DPS
(third) burn and a 0.5 second coast, followed immediately by LEM staging
and a 43Z. 6-second APS (first)burn.
The third DPS burn thrust and propellant weight flow rate profiles are
presented ix,Figures 4-2 and 4-3, respectively. At shutdown, _all of the -
DPS propellants available,except the one percent flight performance reserves,
have been consumed.
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LEM staging and the first/%PS ignition occur simultaneously. It is
assumed that i, 183 pound-seconds of impulse will be delivered to each
stage during LEM staging. Staging characteristics during the first several
minutes after LE_A/[ staging are presented in Figures 4-4 and 4-5. These
data are based upon point n%ass simulations. LEM staging occurs at a
position in the orbit to allow simultaneous tracking of the event by the
Apollo tracking s.hip, Point Arguello, Goldstone, Guayrnas, and White
Sands ground stations. Section Z. 2. Z h) of the Mission Profile Guidelines
suggests that this event be positioned so that at least three ground stations,
with data record capab_//ty, can receive these data.
The spent descent stage is left on an orbit with a period of 91. i
minutes and an inclination of 31. 438 degrees. The perigee altitude is
141.0 nautical miles and *.he apogee altitude is 223. i nautical miles. The
estimated descent stage n-.aximum orbital lifetime is approximately 39
days (see Table 4-Z).
The spacecraft ascent stage propulsion system is c.haracterized by a
vacuum thrust and prope _l!ant %veight flow rate of 3,500 pounds and ii. 45
pounds per second, respectively. The duration of the first APS burn was
chosen so that all the propellants available, except the one percent flight
performance reserves, are consumed over the three suggested burns.
This resulted in a 31.4-second decrease from the duration suggested in
Section 2.2.2j).
As in the second DPS burn, most of the AV available from these
burns {approximately -=,000 feet per second) is dissipated out of the orbit
plane. /% constant inertial attitude is held (see Yig_/re 5-41) from the RCS
ullage maneuver to three seconds after the APS ignitio_ The ascent stage
is then rolled at a constant rate of 0. 0293 degree per second. The 3-second
delay in the maneuver is to allow for hardware clearance. The above atti-
tude and roll rate place the LEM ascent stage, at APS shutdown, on an
orbit with a period of 92.0 minutes and an inclination of 31. 375 degrees.
The orbital perigee altitude is _35.4 nautical miles and the apogee altitude
is 275.8 nautical miles.
4. 10 ORBITAL COAST TO SECOND APS BURN
The suggested 20-minute time duration of this coast to within -+ Z
minutes is included in 2. Z. 2 i) of the Mission Profile Guidelines. Using
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Page 39
this coast time, it becomes necessary to locate Apollo tracking ship No. 3
off the western coast of Africa. The primary function of this ship is to
provide general tracking information, data recording and ground command
capabilities in support of the second A.PS pre-burn and burn events.
No attempt has been made to simulate various I_CS tests expected to
be performed during this coast; however, a certain portion of the RCS
propellants available have been allocated for this phase of the mission
(see Table 5-5). The spacecraft is being tracked approximately 8 minutes
during this coast.
After approximately 14 minutes and 40 seconds of this coast have
elapsed, a I2-second spacecraft orientation maneuver is initiated to achieve,
and hold; the desired pre-burn attitude (see Figure 5-48).
4. il SECOND APS BUR_N
A 3-second RCS ullage maneuver precedes the second APS ignition.
This event is initiated approximately 4 minutes after the Apollo tracking
ship begins tracking the spacecraft.
The inertia/attitude (see Figure 5-48) is held constant during the
5-second burn. This attitude was determined so as to decrease the orbital
perigee altitude. This is done to decrease the expected orbital lifetime of
the ascent stage. The post-burn perigee altitude is t15.8 nautical miles
and the apogee altitude is Z68.3 nautical rni/es. This orbit has a period of
91.5 minutes and an inclination of 31. 2718 degrees.
4.12 ORBITAL COLD-SOAK TO THIRD APS BLq%N
Ten seconds after the second APS burn, an i8-second spacecraft
rnaneuver is performed to align the +X-axis (yaw axis) normal to the
ecliptic and the + Z-axis (roll axis) toward the sun. This attitude is held
( + 5 degrees deadband) for approximately 2 hours and 30 minutes. This
coast duration is slightly shorter than that suggested in the lV[ission Profile
Guidelines. The in-orbit position of the third APS burn achieved by the
shortened coast allowed the burn to occur over MSFN stations with the
necessary ground command and data record equipment. After the attitude-
hold coast, an S-second spacecraft orientation maneuver to the desired
pre-burn attitude is initiated (see Figure 5-57).
In the simulation, the spacecraft attitude drifted approximately 0. i
degree, with no maneuvers, duringthe orbital cold-soak (see Figure 5-53).
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4. t3 THIRD APS BUR_
Two minutes after the Point Arguello tracking site acquires the
spacecraft, a 3-second RCS ullage maneuver is per£ormed and is immediately
followed by a S-second APS burn. All the available propellants are con-
surned By the APS during this burn, with the excep_on of a one percent
flight perforzna, nce reserve.
The initial inertial pitch attitude held during this APS burn (see
Figure 5-57) was chosen to further decrease the orbit perigee altitude.
A't _ shutdown, the spent LEM ascent stage is on an orbit charac-
terized by a perigee altitude of 109.7 nautical miles and an apogee altitude
of 284. 8 nautical nliles. The orbital period is 91.7 _:_utes. The estimated
ascent stage n_drnurn orbital lifetime is approxlu-_a:ely 2"/ days (see
Table 4-2).
4. 14 FINAL OR.BITAL COAST
The objectives of the mission are essentially completed at this point
in the mission; however, if the capability exists, additional tests of the
RCS system _ be performed during this coast. Again, no attempt has
been made to sL-nulate these various tests in this prcf_.ie. A certain portion
of the R.CS propellants available have been allocated for this phase of the
mission. An a_owa_uce of approximately 4.5 hours is shown in this profile.
4. 15 ORBITAL LIFETD_[E ESTIMATES
The approximate orbital lifetimes of various I_3,'-I configurations
have been es_./rnated and are presented herein in the form of days to impact.
The three Basic spacecraft cor_igurations considered are outlined
below:
Confi_,ura._ion I - _ ,
The LEM-1 on the nominal orbit after separation from the
S-IVB/SLA, But Before any major propulsion tests.
Configuration 2
The spent _LEA{-I descent stage on the nominal orbit at
the instan_ of spacecraft separation.
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Configuration 3
The spent LEIV[-I ascent stage on the nominal orbit after
the third A.PS burn has been accomplished.
BaLlistic coefficients, W/CDA (weight divided by the orbital drag
coefficientand the frontal area), were calculated for each of the configura-
tions. An orbital drag coefficient of 2. 0 has been assumed. Various views
of each configuration were studied in order to arrive at the minimum and
the moaximurn frontal area that each configuration could exhibit normal to
the velocity vector. Table 4-I. presents the results of this analysis.
Table 4-1. Ballistic Coefficients
W/CDA
Confi_--uration/Area Weight (Ib) Frontal Area (ft2) (Ib/ft2)
!/* 32,540 200 81
Z /minimum 4, 795 80 30
2/maximum 4, 795 200 IZ
3/minimum 5, Z89 iZ5 Zt
3/maximum 5,289 190 14
*The frontal area of the LE_[-I spacecraft does not appreciably chan_e ".'hen
studied from various views; therefore, only one frontal area is given.
These ballistic coefficients, along with the applicable orbital characteristics
from the nominal trajectory and from Reference 16 were used to calculate
the orbital lifetime estirr_.ates presented in Table 4-Z.
Table 4-Z. Orbital Lifetime Estimates
Con_q._ation
t
2
3
Orbital Life'd.rne (days)Minimum .V.a xirnu__..
6 (See Table 4-1. )
16 39
18 Z7
3300-H007-RC000
Page 4Z
It should be noted that even if the orbital lifetime estimates presented
above are in error by so much as I O0 percent, the), will still fall well within
the 3-month limit for orbital lifetime suggested in the Mission Profile
Guidelines section of this report. Furt_hermore, l_eference 19 states that
1'special preventive measures are not required for the LEM on Ap011o
Mission SA-ZO6A"
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5. NO**_INAL TRAJECTORY DATA
This section contains trajectory parameter histories describing and
illustrating the nominal mission profile. These data, presented in tabular
and graphic forms, are based upon the trajectory printout data in Volume
II of this document." • -
5. i MISSION PROFILE •DATA
• The time sequence of events for the entire mission is shown in
Table 5-I. Figure 5-I presents the earth ground track for the entire
mission. Figures 5-Z and 5-3 present the earth ground track for the two
propulsion system tests that occur over the United States. Orbital
characteristics for the spacecraft coast phases are presented in Table 5-Z.
Earth shadow information (daylight - darkness) is illustrated in
Figure 5-I and presented in tabu/at form in Table 5-3. A time history of
the spacecraft body attitude rates is presented in Table 5-4.
Table 5-5 presents the LEM-I_CS propellant expenditures based on
the information from Reference 6, using the criteria established in Table
3-5.
5. Z TRAJECTORY PHASE DATA
Discrete events summaries and time history illustrations of the
launch vehicle and .spacecraft position, motion, and attitude are presented
for each of the fourteen major phases of'the mission as follows:
TableMission Phase
Saturn IB Ascent to Orbit 5-6
S-IVB/SLA/LEM Orbital Coast 5-7
Spacecraft Separation 5-8
Orbital Cold-Soak to First
DPS Burn 5-9
First DPS Burn 5-I0
Orbital Coast to Second
DPS Burn 5-I i
Second DPS Burn 5-1Z
Orbital Coast to FITH Abort "Test 5- i 3
FITH Abort Test* . 5-14
Orbital Coast to Second A/_S Burn 5-15
Figures
5-4 through 5-1 1
5-IZ through 5-i5
5-16 through 5-20
5-2! through 5-Z4
5-25 through 5-29
5-30 through 5-31
5-3Z through 5-36
5-37 through 5-38
5-39 through 5-43
5-44 through 5-45
The FITH Abort Test phase consists of the third DPS burn, LE:/[
staging, and the first APS burn.
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Page 44
Mission Phase
Second APS Burn
Orbital Cold-Soak to Third
• APS Burn
Third APS Burn
Final Orbital Coast
Table
5-t6
5,17.
5.18
5-19
Figures
5-46 through 5-50
5-51through 5-54
5-55through 5-59
5-60through 5-61
The attitude angles presented in the figures above are referenced to
a launch-centered inertial coordinate system and an earth-referenced
rotating system. These coordinate systems and the spacecraft axis system
are illustrated in Figure 3-7.
3300-H007-RC000
Page 45
Table 5-I. Time Sequence of Events
I
Pha s e Event
Saturn IB Ascent to Orbit
Liftoff/Begin Vertical KisePitch-Over/_Initiate Gravity Turn
End Gravity Turn• S-IB Inboard En=c_Lnes Shutdown
S-IB Outboard EngLues Shutdown/Coast
S-IVB Ign/t_onJettison Thern_olag, Ullage Cases, and
Dummy CSMS-IVB Shutdown into Elliptical Earth Orbit
S-IVB/SLA/LEM Orbit-al Coast
Start of Orbit_ Coast
Maneuver to A/i_n S-IVB X-Axis AlongOrbit Path
S-IVB X-Axis A!i_'-ed Along Orbit Path
Carnarvon Trac___-_ Acquisition
SLA Petal De_loy_e_t
Spacecraft Separa:!on
SLA Petal Deplo_ent
LEM Separa:ion/KCS IgnitionRCS Shutdown
LEM LandLug C_ar Deployment
Orbital Cold Soak to First DPS Burn
LEM Landing Gear Deployment
Maneuver to Aii_-n :._AI +Z-Axis TowardThe Sun
Maneuver to Keq,_red Pre-Buf'n.Inertial A_!.'__de
Carnarvon Traci_.-n_ AcquisitionRCS Ullage .k'aneuver
First DPS Burn
RCS Ullage ManeuverFirst DPS I_r!'/onDPS Shutdown
Orbital Coast to Second DPS Burn
DPS Shutdown
Maneuver to Kequ!red Pre-Burn Inertial• Attitude
Point Arguello Tracking Acquisi¼ionR.CS Ullage A.'aneuver
Time from Liftoff
{hr:min:sec)
0:00:00. O00:00:10. O00:0Z:18. O0O:OZ:ZO. Z50:02:26. 25
0:02:3t. 75
0:02:41.750:09:59.85
O:09: 59.85
O: 10:09.850:10:50.64
0:53:46.26
0:55:46. 26
0:55:46.26
0:55:48.26
0:56:00.26
0:56:08.26
0:56:08.26
0:56:38.26
3:54:28.26
3:59:27.93
4:Or:Z7.93
4:01:27.934:01:35.934:02:07.93
4:02:07.93
4:30:07.934:33:28.244:35:28.24
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3300-H007-RC000
Page 46
Time Sequence of Events (Continued)
Phase Event
Second DPS Burn
RCS Ullage ManeuverSecond DPS IgnitionDInS Shutdown
Orbital Coast to FITH Abort Test
DPS Shutdown
Maneuver to Required Pre-BurnInertia/Attitude
Point Arguelio Tracking AcquisitionRCS Ullage Maneuver
FITH Abort Tect
RCS Ullage ManeuverThird DPS IgnitionDPS Shutdown / Coa st
LEM Staging/First APS IgnitionAPS Shutdown
Orbital Coast to Second APS Burn
APS Shutdown
Maneuver to Required Pre-BurnInertial Attitude
Ship No. 3 Tracking AcquisitionRCS Ullage Maneuver
Second APS Burn
RCS Ullage Maneuver
Second APS IgnitionAPS Shutdown
Orbital Cold-Soak to Third APS Burn
APS Shutdown
Maneuver to Align +Z-Axis TowardThe Sun
"- Maneuver to Required Pre-BurnInertial Attitude
Point Arguello Tracking AcquisitionRCS Ullage Maneuver
Third APS Burn
RCS Ullage ManeuverThird APS IgnitionAPS Shutdown
• Final Orbital Coast
APS ShutdownEnd of Mi||ion Profile
Time from Liftof£
• (hr:min: sec) ,
4:35:28.24
4:35:36.24
4:47:46. 24
4:47:46.24
6:06:06.246:09:25.776:12:45.77
6:12:45.77
6:12:53.77
6:t3:20. 776:13:21.27
6:20:24. 87
6:20:24. 87
6:35:04. 87
6:36:29.256:40:24. 87
6:40:24.876:40:27.87
6:40:32.87
6:40:32.87
6:40:42. 87
9:10:32.87 _
9:22:53.339:24:53.33
9:24:53.339:24:56.33
9:25:01.33
9:25:0t. 33t4:00:00.00
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Table 5-3. Earth Shadow Data
Entrance IntoEarth's Shadow
(Time From Liftoff)
Hrs IV[in
0 38
2 6
3 34
5 3
6 35
8 6
9 37
tt 9
i2 39
Exit From
Earth's Shadow(Time From Liftoff)
Hrs Min
1. 14
2 43
4 ii
5 40
7 it
8 43
i0 14
It 45
i3 17
Time InEarth' s Shadow
MLn
36
37
37
37
36
37
37
36
38
. .
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3300-H007-EC000
" Page 53
Spacecraft Body Attitude Rate History
Time from Liftof£
_nr:min:sec)
Spacecraft Separation0:55:48.26 O.
:Orbital Cold-Soak to First DPS
•0:56:08.26 • O.0:56:38.26 O.0:56:45.060:56:56.6t0:57:3t. 883:54:28.263:54:32. 1.63:54:41.903:55:09.78
Pitch Rate Yaw Rate Ro].].Rate
(de_/sec_ (deg/sec) (de_/sec)
B11rn
00
0.00.00.00;00.00.00.0
Orbital Soak to Second DPS Burn
4:02:07.93. O. 04:30:07.93 O. 04:30:08.10 O. 04:30:16.97 O. 04:30:46.5t O. 0
-iSecond D!=S Burn
4:35:36.24 O. 04:47:46. 24 O. 0
0,0
0.0-5.0
0.0-5.0
0.05.0
0.0-5.0
0.0
0.05.00.05.00.0
0.00.0
+
0.0
0, 00.0
t0.0
0.00.00.0
10.0
0.00.0
0.00.0
tO.O0.00.0
_. 0t66O. 0166
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Orbital Coast4:47:46. 246:06:06. 246:06:06.926:06:07.506:06:13.23
FrrH Abort Test6:12:45.776:!3:24. 27
6:20:24. 87
to FITH Abort Test0.00.0
-iO.O
0.0
0.0
Orbital Coast6:20:24. 876:35:04. 876:35:06.946:35:15.936:35:16.76
O.O0.00.0
to Second APS Burn0.00.00.00.00.0
Orbital Cold-Soak to Third APS Burn6:40:32. 87 O. 06:40:42.87 O. 06:.40:50, 72 O. 0 ,6:40:55.83 + 0.06:41:0[. 13 O. 0
9:30:32. 87 O. 09:10:34. 80 O. 09:10:38. 64 O. 09:t0:41.30 O. 0
0.0-5.0
0.05.00.0
0.00.00.0
0.0-5.00.05.00.0
0.05.0
0.0-5.0 ....
0,0-5.0
0.0-5.0
0.0
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0.0O. 0293O. 0293
Oo 0 -
0.0-t0.0
0.00.0
0.00.0
10.0
0.00.0
0.0-10.0
0.00.0
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Table 5- 5. ZiEIVi-RCS Propellant Expenditures*
Maneuver
Spacecraft Separation
Ullages Preceding
I) DPS Operation
2) A'I_S Operation
Attitude Holds (+5 deg Deadband)
I} During LEM Coast
Z) During Ascent Stage Coast
Attitude Holds [_0.3 deg Deadband)
I) During DPS Burns
2) During APS Burns
3) During FITH Staging
Three Axis Attitude Orientation
t) L_.-M
2) Ascent Stage
RCS Tests***
i) Coast Between First and Second DPS Burns
Z} Coast Between Second and Third DPS Burns
3) Coast Between First and Second APS Burns
-4): Coast After Third APS Burn
Total =
Usable Propellant Remaining**** =
RCS
PropellantExpenditure
(Ib)**
15.00
30. O0
7.50
0.76
3.90
31.50
64. 18
10. O0
70.40
8.20
18.84
91.22
44. 89
44. 89
441.28
134. 72
*No allowances were made for RCS contingency operations.**Based on the criteria from Table 3-5 and from mission profile.
***Reference 6.
-***_Based on an RCS usable propellant loading of 576 pounds. "
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3300-II007-RC000
Page 126
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3300-II007-I_C000
Page i 27
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3300-H007-IIC000
Page i28
6. TitACI(IiNG _AND COlvII'ZUN...CA2[IOI,_S"" r " _ D./kTYk
' i
Spacecraft visibility periods for the MSFN stations presented in
"Table 3-6 are listed in Table 6-i. Table 6-2 presents the intervals of the
mission which are not seen by any of the tracking stations (comrnunlcations
void). The surface tracking coverage, during the ascent to orbit, space-
craft separation and staging, and all the 7kPS and DPS burns, is shown in
_'igure 6-I. Spacecraft visiblity is defined as a tracking elevation angle
greater than 5. 0 degrees as measured from the station local horizontal.
Volume Ill presents detailed tracking tin]e history date. for the ground
stations available for operation on this mission. These data consist of
range, range rate, azin_uth angle, elevation angle, and two spacecraft-to-
radar look angles, and are presented as a function of time for each of the
ground stations. Significant events ar'e noted in this data.
t
Table 6- i. Surface Tracking Coverage
3300-H007-RC000
Page IZ9
Trackin K Station
Grand Turk
Cape Kennedy
Grand Bahama
San Salvador4
B emnud a
Grand Canary
Carnarvon
Ship No. i
Guayma s Mex.
White Sands
Texas
Cape Kem_edy
San Salvador
Grand Bahama
B ermuda
Grand Canary
Carnarvon
Ship No. I
Hawaii
Ship No. 2
l_t. Arguello
Ooldstone
Ouayma s Mex.
White Sands
Te.x_4, s
Cape Kennedy
Grand 13ahanna
San Salvadoz'
Grand Turk
}3ern,_uda
Ascension Is.
Ship No. 3
Acquisition
of Signal-Time from Liftoff
__(hr :_nin:sec)
Loss of Signal-Time fron_ Liftoff
__ ID£2." o }__
0:05: 47. 251 0:07:36. 155
0:00:19. 332 0:07:56. 850
0:01:55. 066 0:08:i6,833
0:03:10. 837 0:08:34. 194
0:05:57. 79Z 0:iI:51. 464
0:I8:il. 446 0:Z2:48, 459
0:53:46. 255 0:57:26. 467
0:52:48. 850 0:58:i4,320
1:30:02.8tI I:34:32. 386
1:3i:36. £6i i:35:31.470
i:33:24,034 I:37.57.064
i:36:56, I04 1:40:57, 147
I:39:55. 908 i:40:i5. 058
i:37:45. 816 i:41:04,236
i:40:I0. 280 I:44:47.569
I:5Z:49. 890 I:53:16. 695
2:Z6:49. 787 2:3I:07.915
g:Z5:49.235 Z:31:i5.095
2:5Z:51. 470 2:55:22.40Z
3:00:I 3. 931 3:04:43. 612
3:01:Z9.179 3:05:23. 050
3:0Z:I 8. 832 3:06:09.43i
3:03:08. 545 3:07:i6,433
3:04:03. 940 3:08:37. I38
3:06:3i. 995 3:10:5i. 530
3:09:56. 585 3:14:i I. 884
3:I0:4i.574 3:I4:36. I87
3:12:06.545 3:15:I3,213
3:13:51.382 3:15:18. 67Z
3:1 3:09. 569 3:17:23. 260
3:29:52. 307 3:32:17. 357
3:27:15. 925 3:32:18. 010
VisiY_ilityDuration
__(mi : oeK
i:48. 904
7:37. 5i8
6:2i. 767
5:23. 357
5:53. 672
4:37. 0i3
3:40. 2If
5:25. 470
4:29. 575
3:55. 3i0
4:33.030
4:0I. 043
0:19. i50
3:i8.4i9
4: 37. 289
0:26. 805
4:I8, 128
5:Z5. 859
g: 30.9 32.
4:29. 681
3:53. 87i
3:50. 599
4:07. 888
4:33. 199
4:19. 534
4:15. Z99
3:54. 613
3:06. 668
1:2.7.290
4:13. 691
2:2.5. 050
5:02. 085
Tabl;6-I.
3300-Ii007-I_C000
Page i 30
Surface Tracking Coverage (Continued)
Trackin_ Station
Pretoria
Ship No. i
C3.rn&rvon
I-lawaii
Ship No. Z
Pt. Arguello
Goldstone
Guayma s Mex.
White Sands
Texas
Cape Kennedy
Grand B aha:r_
San Salvador
Grand Turk
Bern_uda
Antigua
Ascension Is.
Ship No. 3
Pretoria
Ship No. I
Carnarvon
Hawaii
Ship No. 2
Pt. Arguello
Goldstone
Oue, yn_a s Mex.
White Sands
Texas
Cape Kennedy
Grand Bahama
San Salvador
(]rand Turk
I3 e rmu da
Acquisitionof Signal-
Time from Liftoff
___hr:min: s ec)
3:39:53.8i6
3:59:03. 035
3:59:Z7.93Z
4:24:13. 836
4:32:39. 315
4:33:28,239
4:34:17. 826
4:35:5Z. 0Zi
4:36:30. 704
4:39:03. 227
4:42:25. 312
4:43:08. 439
4:44:Z4. ii8
4:45:45. 770
4:45:48. 858
4:49:07. 508
5:03:Z9. 420
5:00:23.808
5:13:i0. 388
5:32:i7. 219
5:3Z:50. 370
5:59:Z3. 175
6:08:26. 660
6:09:25. 772
6:i0:i8. 916
6:ii:4i. 628
6:12:31.502
6ii5:06. 373
6:18:3Z. 619
6:i9:i4. 850
6:20:31. 369
6:2i:57.695
6:22:00. 441
Loss of Signal-Time from Liftoff
___Jhr :n_in:sec____
3:45:05. 434
4:03:38. 476
4:04:45. Z57
4:29:52. 525
4:38:44. 995
4:39:35.237
4:40:23. 267
4:4I:28.259
4:42:44. 528
4:45:06. 936
4:48:29. 985
4:48:54. 293
4:49 38, _" _: _3t,
4:50:07. 605
4:5i:47. Z03
4:5i:54.875
5:06:Z0.4Z5
5:06:49.171
5:20:16. 475
5:40:24.4iI
5:41:i6. 890
6:06:27. 794
6:14:55. 900
6:15:40. I07
6:i6:Z6. i78
6:i7:39. 046
6:i8:48. 977
6:21:11.454
6:24:29. 362
6:24:52. 706
6:25:34.84i
6:25:59. 271
6:27:49.61Z
VisibilityDuration
e c)_5:11. 618
4:35.44i
5:17. 325
5:38. 690
6:05. 680
6:06. 999
6:05. 441
5:36. 238
6:i3. 823
6:03. 710
6:04. 673
5:45. 854
4:2I. 835
5:58:345
Zi47. 368
Z:51. 005
6:25. 364
7:06. 087
8:07. IgI
8:Z6. 519
7:04. 619
6:29. 240
6:14. 335
6:07. 262
5:57. 418
6:i7. 474
6:05.08i
5:56. 743
5:37. 856
5:03. 471
4:01. 575
5:49. 17 i
Table 6- i.
3300-}i007-Z_C000Page 131
Surface Tracking Coverage (Continued)
Trackin_ Station
AntiguaAscension Is.
Ship INo. 3Pretoria
Ship No, iCarnarvon
Hawaii
Ship No. 2
Pt. ArguelloGolds tone
Guayma s Mex.
White _ -'oanus
Te×_,s
Cape Kennedy
Grand Dahama
San Salvador
B ernnuda
Grand Turk
Antigua
Ship No. 3
Ascension Is.
Pretoria
Ship No. i
Carnarvon
Gua.n'l
Hawaii
Pt. Arguello
Ship No. 2
Goldstone
G,uaymzs Mex_
Y,rhit e Sands
T C Y_. S
Cape Kennedy
Acquisition
of Signal-Tinae from Liftoff
___(hr:n_in:sec)
6:25:39. 36Z
6:39:48. 361
6:36:29. 247
6:49:02. 437
7:08:26. 337
7:09:02. 751
7:36:32. 552
7:45:32. 055
7:46:34. 339
7:47:27. Z57
7:48:43.80Z
7:49:31. 813
7:52:03. 342
7:55:19.0Zi
7:55:58.510
7:57:01. 447
7:59:46.02.4
7:58:02. 963
8:00:29. Z39
8:13:07. 508
8:i3:33.696
8:25:00. 845
8:46:11. 636
8:46:09. 943
8:59:58.4gi
9:i3:55. 813
9:22:53. 326
9:22:02. 780
9:23:47. 814
9:25:06. 053
9:25:59. 381
9:28:25.'464
9:32:19. Z86
Loss of Signal-Tin-le frozn Liftoff
....(br :n_in:sec)___
6:27:26. 348
6:42:17. 625
6:43:03. 953
6:57:07. 692
7:17:35. 094
7:i8:gZ. 595
7:42:44. 092
7:50:58. 900
7:51:41. 945
7:52:25. 581
7:53:47.183
7:54:51. 806
7:57:g5. 773
8:00:50.06i
8:01:31. 662
8:02:37. 268
8:02:Z2. 799
8:03:36.5Z7
8:06:16. 884
8:19:22. 094
8:ZI:28.000
8:34:11. 435
8:52:16. 953
8:54:12.567
9:05:2,4.385
9:i8:28.917
9:27:50. 943
9:2,7:26.010
9:28:27. 798
9:30:38. 120
9:31:03. 035
9:33:43,115
9:36:19.02,8
VisibilityDuration
1:46. 986
2o4Z:29.
6:34. 705
8:05. 255
9:08. 757
9:19:844
6:il, 54i
5:26. 845
5:07. 606
4:58 324
5:03 381
5:19 992
5:22 4:30
5:3i 04i
5:33 !52
5:35 821
2:36 775
5:33 564
5:4/ 645
6:14. 586
7:54. 304
9:10. 591
6:05. 3i7
8:02. 62.4
5:25:964
4:33. i04
4_57. 617
5:23. 230
4:39. 984
5:32. 066
5:03. 655
5:17. 651
3:59. 752
Table 6- i.
3300-H007-R C000
Page i 32
Surface Tracking Coverage (Continued)
Tracking. Station
Grand Bahama
San Salvador
Grand Turk
Antigua
Ascension Is.
Pretoraa
Gua_n
I-lawaii
Pt. Arguello
Goldstone
Ship No. 2
White Sands
Guayma _ Efex.
Pretoria
Hawaii
Pretoria
Acquisition
of Signal-Time from Liftoff
____(hr: n_in: s ec _.___
9:32:44. lIZ
9:33:38.4:4:8
9:34:27.8Z7
9:37:02.60Z
9:51: 06. 807
10:02:02. 787
10:36:20. 887
10:52:35. 449
11:00:03. 130
11:01:43. 499
10:58:44. 630
11:0,±:51, 811
ii:01:53, ZZ3
1i:39:21. 014
IZlZ8:41.494
13:I6:3Z. Z96
Loss of Signal-Time from Liftoff
_____Jhl':Ir_in:s e c)__.__
9:37:17. 198
9:38:38. 312
9:39:56. 449
9:42: 37. 024
9:57:46. 241
10:11:38
I0:4Z:33
10:54:20
11:03:26
11:03:09.
ii:03:56.
I!:05:01.
li:07:Og.
11:48:59.
IZ:3Z:4Z.
13:Z5:55.
407
092
597
429
53Z
334
Z81
201
78Z
41Z
831
Visibility
Duration
ec)
4: 33. 085
4:59. 865
5:28. 62Z
5:34. 4ZZ
6: 39. 434
9: 35. 620
6: iZ. 205
1:45. 148
3:23. Z99
i:Z6. 033
5:11.705
0:03. 471
5:08. 978
9:38. 768
4:00.9i8
9:23. 535
Table 6-2.
3300-l-1007-FIG000
Page 133
Com_nunications Void Intervals
Void 13egins ,-Tinge fron_ Liftoff
___(hr :mi n:sec)
0:00:00. 000
0:if:51
0:Z2:48.
0:58:14.
1:44:47.
1:53':16,
2:31:15.
2:55:22.
3:17:23.
3:32:18.
3:45:05.
4:04:45.
4:29:52,
4:51:54.
5:06:49
5:20:16
5:41:16
6:06:27
6:27:26.
6:43:03
6:57:07.
7:18:2Z.
7:42:44.
8:06:16.
8:34:11
8:54:12.
9:05:24.
9:18:Z8.
9:42:37
9:57:46
I0:II:38
•464
459
320
569
695
O95
402
Z60
OiO
434
257
525
875
17i
475
89O
794
348
953
692
595
09Z
884
435
567
385
9i7
024
241
407
Void Ends -Timc fron_ Liftoff
___(hr:ITlin:Sec )
0:00:i9. 33Z
O:i8:il, 446
0:53:46. 255
i:30:OZ. 8il
1:52:49. 890
2:26:49. 787
2:5Z:51. 470
3:00:13.93i
3:Z9:52. 307
3:39:53, 816
3:59:03. 035
4:24:1 3. 836
4:3Z:39. 315
5:03:29. 420
5:13:i0. 388
5:3Z:17. 2i9
5:59:Z3. 175
6:08:26. 660
6:_:48. 36i
6:49:0Z. 437
?:08:26. 337
?:36:32. 552
7:45:32. 055
8:13:07. 508
8:46:11.636
8:59:58. 421
9:i3:55. 813
9:22:53. 326
9:51:0"o.807
i0:02:02. 787
I0:36:Z0. 887
VoidDuration
0:19. 332
6:i9. 982
30:57. 796
31:48. 491
8:02. 321
33:33, 092
21:36. 375
4:51. 529
12:29. 047
7:35. 806
13:57.60i
19:28. 5"79
2:46. 790
Ii:34, 545
6:21, Z!7
iZ:O0,744
i8:06. 285
1:58. 866
12:Z2 0i3
5:58 484
li:i8 645
18:09 95'7
2:47 963
6:50 624
12:00 20i
5:45. 854
8:3i. 428
4:Z4. 409
8:Z9. 783
4:i6. 546
24:42. 480
,t
Table 6- 2.
Vo!a Begins -
Time from Liftoff
___hr:min:sec)
I0:4Z:33. 092
i0:54:20.597
Ii:03:56. 334
li:07:0Z. 201
i i:48:59.78Z
i2:32:42, 4i2
i 3:2'5:55. 831
3300_l10nv n_r_r_n
Page
Communications Void Intervals (Continued)
V oF_-_277_2,i_-Time from Liftoff
__(hr :n_in: sec)
I0:5Z:35. 449
Ii:00:03. 130
ii:04:57. 81i
ii:39:2i. 0i4
i2:Z8:4i. 494
13:i6:32.296
14:00:00. 000
134
Void
Duration.
.(m_.n: __ c)_.
i0:02. 357
5:42.. 533
I:0i. 477
3Z:I8. 813
i9:4i. 7i2
43:49. 884
34:04. I69
Total Void Time
Percent of Mission
512:07. 729
61.44
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3300-H007-RC000
Page i 37
7. SUMMAP_Y OF TECHNICAL AC}ILEVEMEiNT
This report contains no innovations or improvements involving new
technology, approaches, methods, or patentable ideas as defined in the
co._itract's "New Technology gnd Property P_ights in Inventions' clause..
• APPENDIX
OPEN-LOOP MANEUVER LOGIC
3300-H007-R.C000
Page 138
C---
The purpose of this appendix is to indicate the open-loop type
logic usedto simulate the spacecraft attitude change maneuvers in
inertial space. This logic is sin_i]ar to that expected to be used by the
Apollo spacecraft. The reorientation will consist of a roll maneuver
followed by a pitch or yaw n_aneuver and, if necessary, another roll
maneuver. The naagnitude and direction of the maneuvers are based
upon Euler angles measured from the current attitude orientation to
the desired attitude orientation.
Figure(A-l) shows the Euler angle transfor:_nation required to
change from one inertial attitude to another. These Euler angles are
computed using the knowledge of the unit vectors which describe current
and desired orientation of the spacecraft roll, yaw, and pitch axes. The
components of these Unit vectors are naeasured in the Greenwich inertial
coordinate system at the time of launch. Also calculated is the time•
required to perform the maneuver using the given spacecraft rotational
rates.
The Euler angles are defined as•follows:
the azimuth angle 1_aeasured in the plane formed
by the Y and Z body axes measured from the +Z
body axis to the vector N in the direction of the
-Y body axis.
0 = the polar angle measured from the initial roll
axis (X o) to the final roll axis (Xf).
_J = the azimuth angle n]easured in the new plane
formed by the Y and Z body axes after theand O rotations and is measured fron_ N to [he
final Z body axis (Zf).
The logic uses the Euler angles to con_pute the maneuver angles,
O"roll(1)' _yaw' _pitch' and C_.ro]l(2). The first rnaneuver is a roll
to the closest pitch or yaw axis (Ic_ roll(1)I __ 45 degrees). The second
maneuver is a pitch or yaw of ÷-0 degrees. The third maneuver,
is the final roll and is dependent on the first two maneuvers.e_ roll(2)'
K
I ¸ .i
°'_'i
i
J
3300-H007-RC000
Page i 39
The maneuver angles are defined as follows:
aroll(l ) = the first roll the spacecraft has to performto reorient its attitude,
v pltcl_or
ayawI = the secon,.l nuaneuver the spacecraft has to
perforn_ to reorient its attitude (by definition
one of the two angles is always zero).
_roll(2= the third maneuver (second roll) the spacecraft
has to perform to reorient its attitude.
A
XoA
I _¥.
Figure A-I. Eu[er Angle Transformation
i
/:
i
i,
0
o
4.
1
,
o
,
e
I0.
II.
12.
13.
14.
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Page 140
I_E FEI_EN CES
"Mission l%equireInents for Apo!lo Spacecraft DevelopIuentMission 206A (LEM-I)';MSC Internal Note No. 65-PL-I
(Revision A), from Systems Engineering Division, dated Ii
May 1965.
"Comments o11Revised Mission Requirements for SA-206A",
from FM/Chief, Mission Planning and Analysis Division, dated17 June 1965.
"Saturn IB Control AVeights Analyses", Chrysler Corporation
TB-AE-65-117, from Advance Engineering Branch, dated
1,-February 1965.
"Apollo Mission 206A-Preliminary Mission Profile (U)-Volume I",TRW/STL 3300-H001-RC000, R. K. Petersburg, dated 31 March
1965.
"Flight Mechanics, Dynamics, Guidance and Control Panel Inter-face Control Docun%ent-Saturn 11% SA-206 (BP-30, LEM-I) (U)",
MSFC 80M90206, no date.
"LEM- i Preliminary Mission Capability Report -Mission SA-206A
(Draft)" , GAEC LED-540-38, from LEM Mission Analysis Group,dated I5 June t965.
"Data for SA-Z06A PRT", from ATSO to FAB/C. R. Huss, dated
23 June i965.
"Detailed Test Plan for LEM-1 (First Draft) 'i, GAEC LP7,-6i.t-.3,
fron_ Flight Planning and Analysis, dated 15 February !965.
"Mass Properties D_ta for SA-Z06A, LEM Aio_.e Mission", MSC
P55/M612, from P55/Chief, Design 7Jltegration Branch, dated
ZZ January 1965.
"A Reference Atmosphere for PatrickAFB, Florida", NASA
Technical Note D-595, O. E. Sn_ifll, dated March 1961.
"U. S. Standard Atmosphere, 196Z", U. S. Go.vernment Printing
Office, _Vash._ngton, D. C. , 1962.
"LEM Familiarization Manual", GAEC LIMA 790-1, dated 1"5
Jiffy 1964.
"Aerodynamics Data Manual", North An_erlcan Space a._d Inforn_a-
tion Systems DivisiOn, Vol. AIkM 2-I, Page i. 2.5-I, Revised
I July 1964.
"Minutes of the Tenth Guld_.nce and Perforn_ance Sub-Panel",
Enclosure 10, dated Z0 April 1965.
d
c'T
J
t6.
I7.
I9.
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Page i41
REFERENCES (Continued)
"Apollo Missions and Navigation System Characteristics", NASA-
Apollo Navigation Working Group Technical l'<eport No. 65-AN-I. 0,
dated 5 February i965.
"Lifetime of Near Earth Satellites in Circular or Elliptical Orbits
NASA/MSC, O170 (JCtB:jec), dated 13 September i963, (C).
,I
"A Preliminary Separation Study for SA-200 Tower Jettison with
the CSM Shroud", from FM3/Flight Analysis Branch/MSC, dated
I7 February i965.
"Operational Support Plan for the Apollo 200 Series Missions" I
prepared by the Flight Control Division/MSC, dated April t965.
"Policy Guidance on Orbital Debris", Letter from NASA Head-
quarters/S. C. Phitlips to MSC/W, A. Lee, dated 7 June 1.965.
..... %
APS
CSM
DPS
ECS
EPS
EST
ETR
FITH
GMT
LEM
LES
LMP
MSC
MSFC
MSFN
. RCS
SLA
deg
er
ft
hr
km
ib
rain
n mi
rad
sec
ABBREVIATIONS
Ascent Propulsion System
Con%n-land and Service Module
Descent Propulsion System
Environmental Control System
Electrical Power System
Eastern Standard Time
Eastern Test Range
Fire-ln-The-Hole
Greenwich Mean Time
Lunar Excursion i%4odule
Launch Escape System
LEM IN_ission Progranan%er
Manned Spacecraft Center
Marshall Space Flight Center
Matured Space Flight Net
Reaction Control System
Spacecraft LEM Adapter
degrees
earth equatorial radius
feet
hours
kilometers
pounds
minute s
nautical mile s
radians
seconds
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