C D[- 7 7J - ibiblio Mission SA-206A Spacecraft... · i 3300-hoot-rc-ooo vol. i 1 " _ "c_d[- 7_7j i i apollo mission sa-206a spacecraft pre liaiin._lry reference trajectory (u) i

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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: SPACECRAFT PRELIMINARYI

.... 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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II

Approvedby . V. Stab,efoj_

ManagerManned Spaceflight

Department

Approved byE. A. Ward

........ Ma nagerMission Planning and

Operations, MTCP

_J SYSTE;'_S

" _ 3300-H008-RC000

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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

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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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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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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 .........

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137

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REFERENCES ............................... , , . 140

ABBREVIATIONS . . . eee..e,e..e.ee.eeeeeeeeeeee'e

142

Total Pages: 152

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,.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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ILLUSTRATIONS (Continued)

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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

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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)

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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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• 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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5-18

5-t9

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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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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).

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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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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.

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•,..•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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_.. __i_.-o" Second APS burn: 5 seconds.

"_ = _ '_*_*:;_"_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

3300-H007-RC000

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Event

Weights (Ibl

-1,297=088

435, Z59

II

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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

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0:20 0:40 I:00 1:20

0.2

.,er"-11

O!

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I',--

.<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

Page .17

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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age 18

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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Page 23

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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Page Z6

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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i i•'!i-1200

_ , 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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" ' _ I l °

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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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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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ge 41

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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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

. , v ,.

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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

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3300-H007-EC000

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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

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0.00.0

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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

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0.0-t0.0

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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Page 126

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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.

3300-H007-RC000

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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