3902-H012-RO-000 TRW NOTE NO. 66-FMT-402 PROJECT … Apollo M… · SPACECRAFT REFERENCE TRAJECTORY...

3902-H012-RO-000

TRW NOTE NO. 66-FMT-402

P R O J E C T A P O L L O , ] 2 jggg

A P O L L O MISSION A S - 2 0 6 AS P A C E C R A F T R E F E R E N C E T R A J E C T O R Y

Prepared byTrajectory Design Section

TRW Systems

VOLUME ITRAJECTORY DESCRIPTION

28 FEBRUARY 1966

-CR-1332B3) A P O L L C M I S S I O N AS-206AC B A F T B E F E R E B C E T B A J E C 1 C B Y . V O I O M E

1: TRAJEC10BY DESCRIPTION ( T R W SystemsJ G r o u p ) 92 p

N73-73426

Unclas00/99 17962

Prepared forMISSION PLANNING AND ANALYSIS DIVISION

NATIONAL AERONAUTICS AND SPACE ADMINISTRATIONMANNED SPACECRAFT CENTER

HOUSTON, TEXAS

3902-H012-RO-000

RKP

TRW NOTE NO. 66-FMT-402

P R O J E C T APOLLO

A P O L L O MISSION AS-206AS P A C E C R A F T R E F E R E N C E T R A J E C T O R Y

28 FEBRUARY 1966

VOLUME ITRAJECTORY DESCRIPTION

Prepared for . -MISSION PLANNING AND ANALYSIS DIVISION

NATIONAL AERONAUTICS AND SPACE ADMINISTRATIONMANNED SPACECRAFT CENTER

HOUSTON, TEXASNAS 9-4810

Approved by:.Co Ro Huss, ChiefFlight AnalysisBranchNASA/MSC

Approved by:.JohhlP. Mayer, ChiefMissipn Planning andAnalysis DivisionNASA/MSC

Prepared by: R. M. DiamondW. B. GrayR. D. ShinkleF. J. SichiT. L. Vogr~^TRW Syste

Approved by:C. V. Stableford^Acting ManagerMission Design DepartmentTRW Systems

Approved by:. W.^Pittman, Manager

' Mission Planning andOperationsTRW Systems

TRWSYSTEMS

FOREWORD

This report, which defines the Spacecraft Reference

Trajectory for Apollo Mission AS-206A, is submitted to

NASA/Manned Spacecraft Center by TRW Systems in par-

tial response to Task MSC/TRW A-21 (Establishment of the

Reference Trajectory for Apollo Mission AS-206A), with«

Amendment Number 1, of the Apollo Mission Trajectory

Control Program (Contract No. NAS 9-4810).

This report is presented in three volumes. Volume

I summarizes the mission requirements, the unclassified

spacecraft input data used in the generation of the mission

profile, and the significant trajectory analyses. Volume I

also contains graphical and tabular time histories of perti-

nent spacecraft attitude, position, motion, separation char-

acteristics, and MSFN coverage data. Volume II of this

report contains the trajectory listing of selected mission

profile parameters, and the confidential spacecraft input

data specifications. Volume III presents detailed time his-

tories of radar range, range rate, elevation, elevation rate,

azimuth, azimuth rate, and four look-angles for the MSFN

ground stations available for operation during this mission.

111

• CONTENTS

Page

1. INTRODUCTION AND SUMMARY . . ................ 1

1. 1 Purpose ..................... . ......... 1

1. 2 Scope .............................. • • • 1

1. 3 Mission Profile Summary ................... 2

2. SPACECRAFT MISSION .REQUIREMENTS. ... ......... 9

2. 1 Spacecraft Test Objectives ................... 9

2. 1. 1 Primary Spacecraft Test Objectives ....... 92. 1. 2 Secondary Spacecraft Test Objectives ...... 9

2. 2 Mission Profile Guidelines ... ............... 9

2. 2. 1 Launch Vehicle Guidelines ........ ...... 92. 2. 2 MSFN Coverage /Flight Control Guidelines ... 92. 2. 3 LEM Attitude Guidelines . . . . . ..... ..... 102. 2. 4 LEM Propulsion Guidelines ............ ". 112. 2. 5 General Mission Guidelines . . .......... . 12

3. SUMMARY OF INPUT DATA . . . . ........ . . . ..... :. 13

3. 1 Saturn IB Launch Vehicle .................... 13

3. 2 Spacecraft (LEM-1) ................. ...... 13

3. 3 MSFN Stations ........................... .13

3. 4 Earth Constants and Conversion Factors ......... 14

3. 4. 1 Earth Constants ..................... 153. 4. 2 Conversion Factors . ................. 15

3. 5 Spacecraft and Attitude Reference CoordinateSystems ............... . ............... 16

3. 5. 1 Spacecraft Coordinate System,. X«, Y^, Zg. . . 16

3. 5. 2 Launch S_ite Inertial Reference System,x r Y r z i . . . . . . . . . . . . . . • • • . . . . . . ' • • 1 6

3. 5. 3 Launch Site Rotating Reference System,XR' YR' ZR

3. 5. 4 Relative Vehicle Coordinate System,XRV YRV ZRV

CONTENTS (Continued)

Page

4. MISSION DESCRIPTION. 25^ ' f -

4. 1 As cent-to-Orbit . 25

4. 2 S-IVB/SLA/LEM Orbital Coast .'.. . . . . . . . . . .26

4. 3 Spacecraft Separation 26

4. 4 Orbital Cold-Soak to First DPS Burn 27

4. 5 First DPS Burn . . . . 27

4. 6 Orbital Coast to Second DPS Burn. . 28

4. 7 Second DPS Burn/FITH Abort Test/First APS Burn. . 28

4. 7. 1 Second DPS Burn 284. 7. 2 FITH Abort Test/First APS Burn . . . . . . . . . 30

4. 8 Orbital Coast to Second APS Burn 30

4. 9 Second APS Burn . . 30

4. 10 Orbital Coast to Third APS Burn. . 32

4.11 Third APS Burn. 32

4. 12 Orbital Cold-Soak to Fourth APS Burn 32

4. 13 Fourth APS Burn 33

4. 14 Final Orbital Coast 33

4. 15 Spacecraft Orbital Lifetime Estimates 33

5. NOMINAL TRAJECTORY DATA 41

5. 1 Mission Profile Data ' . - . . . . 41

5. 2 Trajectory Phase Data 41

6. MSFN COVERAGE DATA ' 67

7. SUMMARY OF TECHNICAL ACHIEVEMENT 79

REFERENCES 81

VI

TABLES.

Page

3-1 MSFN Ground Stations and Capabilities 17: - '3i • • '. j

4-1 Ballistic Coefficients : 34

4-2 . Spacecraft Orbital Lifetime Estimates. 34

5-1 Time Sequence of Events 42

5-2 Orbital Characteristics of the Spacecraft Coast Phases. . 44

5-3 Spacecraft Earth Shadow Data. . . . . . 45

5-4' • Ascerit-to-Orbit; Discrete Events Summary . . . . . . . . . 46

5-5 SLA Petal Deployment and Spacecraft Separation;Discrete Events Summary 47

; ' . . • " - • - . " : . ' ' ' ' • - '

5-6 First DPS Burn; Discrete Events Summary. . . . . . . . . . 47

5-7 Second DPS Burn/FITH Abort Test/First APS Burn;Discrete Events Summary.. ... . . . . 48

5-8 Second APS Burn; Discrete Events Summary. . . . . ; . .-.' 48

5-9 Third APS Burn; Discrete Events Summary 49

5-10 Fourth APS Burn; Discrete Events Summary. . 49

6-1 MSFN Mission Coverage 68

6-2 Communication Void Intervals 74

Vll

- ILLUSTRATIONS

Page

1-1 Mission Profile Summary '. . ' 5'

3-1 Saturn IB Launch Vehicle Outboard Profile. . . . . . . . . . 18

3-2 Spacecraft Reference Dimensions 19

3-3 LEM-1 Outboard Profile . . . . . . . . . . . . . . . . . . . . . . 20

3-4 Australia Apollo Tracking Ship Placement Study/Spacecraft Separation . . . 21

3-5 Australia Apollo Tracking Ship Placement Study/First DPS Burn 21

3-6 Bermuda Apollo Tracking Ship Placement Study 22

3-7 Spacecraft Coordinate System. 23

3-8 Spacecraft Reference Coordinate System 23

3-9 Relative Vehicle Coordinate System 24

4-1 Spacecraft Separation; Relative Velocity andDistance during RCS Thrusting 35

4-2 Spacecraft Separation; Relative Velocity andDistance History for Approximately 100 Minutes . . . . . . 35

4-3 Spacecraft Cold-Soak Attitude Orientation. . . . . . . . . . . 36

4-4 Cold-Soak Attitude Orientation History for Launches,Second Quarter 1967. 37

4-5 First DPS Burn Thrust Profile 38

4-6 Second DPS Burn Thrust Profile 38

4-7 Effect of Assumed IMU Drift on Orbit DimensionsFollowing Second DPS Burn , 39

4-8 FITH Staging; Relative Velocity and DistanceHistory During First APS Burn 39

4-9 FITH Staging; Relative Velocity and DistanceHistory" to 100 Seconds 40

4-10 Effect of Assumed IMU Drift on Orbit DimensionsFollowing Second APS Burn 40

ix

ILLUSTRATIONS (Continued)

Page

5-1 Earth Ground Track; Revolutions One and Two . 50

5-2 Earth Ground Track; Revolution Three 51

5-3 . Earth Ground Track; Revolution Four 52

5-4 Earth Ground Track; Revolutions Five through Nine. ... 53

5-5 As cent-to-Orbit; Altitude, Geodetic Latitude,and Longitude . . . 54

5-6 As cent-to-Orbit; Inertia! Velocity, Inertia!Flight Path Angle, and Inertia! Azimuth . . . 54

5-7 ' As cent-to-Orbit; Dynamic Pressure and MachNumber 55

5-8 As cent-to-Orbit; Launch Site Inertia! AttitudeHistory 55

5-9 Spacecraft Separation; Altitude, Geodetic Latitude,and Longitude 56

5-10 Spacecraft Separation; Inertia! Velocity, Inertia!Flight Path'Angle, and Inertia! Azimuth 56

5-11 Spacecraft Separation; AV History. 57

5-12 First DPS Burn; Altitude, Geodetic Latitude,and Longitude 57

5-13 First DPS Burn; Inertia! Velocity, InertialFlight Path Angle, and Inertial Azimuth. . 58

5-14 First DPS Burn; Launch Site Inertial AttitudeHistory 58

5-15 First DPS Burn; AV History 59

5-16 Second DPS Burn/FITH Staging/First APS Burn;Altitude, Geodetic Latitude, and Longitude 59

5-17 Second DPS Burn/FITH Staging/First APS Burn;Inertial Velocity, Inertial Flight Path Angle, andInertial Azimuth 60

5-18 Second DPS Burn/FITH Staging/First APS Burn;Launch Site Inertial Attitude History £,Q

ILLUSTRATIONS (Continued)

5-19

5-20

5-21 ..

5-22

5-23

5-24

5-25

5-26

5-27

5-28

5-29

6-1

Second DPS Burn/FITH Staging/First APS Burn;AV History

.. - -. • - • i- .Second APS Burn; Altitude, Geodetic Latitude,and Longitude i

Second APS. Burn; Inertial Velocity, Inertia!Flight Path Angle, and Inertial Azimuth

Second APS Burn; Launch Site InertialAttitude History.

Second APS Burn; AV History. . . . . ; . . . . . • . . . . . . .

Third APS Burn; Altitude, Geodetic Latitude," andLongitude

Third APS Burn; Inertial Velocity, Inertial FlightPath Angle, and Inertial Azimuth

Third APS Burn;" AV History . . . . . . . . . . ... ; . . . .

Fourth APS Burn; Altitude, Geodetic Latitude,and Longitude . '-'. . •* . . . . . . .

Fourth APS Burn; Inertial Velocity, InertialFlight Path Angle, and Inertial : Azimuth. . . . . . .'.•". . .

Fourth APS Burn; AV History . . . . . . . . . ; . ' . .

MSFN Coverage Summary

Page

61

61

62

62

63

63

• • • : 64

,- .-•• 64

, • 65

. 65 '

• - 66

76

XI

.. • ABBREVIATIONS

APS Ascent Propulsion System

AGS Abort Guidance Subsystem -., '.

GSM Command and Service Module

DPS Descent Propulsion System

EST Eastern Standard Time

FIT-H Fire-In-The-Hole

GMT Greenwich Mean Time

IMU Inertial Measurement Unit

LEM Lunar Excursion Module

LES Launch Escape Subsystem

LMP LEM Mission Programmer

NASA National Aeronautics and Space Administration

MAX Maximum

MCC Mission Control Center

MIN Minimum

MSC Manned Spacecraft Center

MSFC Marshall Space Flight Center

MSFN Manned Space Flight Network

PCM Pulse Code Modulation

PNGCS Primary Navigation and Guidance Control System

RCS Reaction Control System

SLA Spacecraft LEM Adapter

SLR Radar Slant Range

UHF Ultra High Frequency

VHF Very High Frequency

XI11

ABBREVIATIONS (Continued)

deg Degrees

er Earth Equatorial Radius

ft Feet

hr Hours

km Kilometers

lb Pounds

min Minutes

n mi Nautical Miles

rad Radians

sec Seconds

xiv

1. INTRODUCTION AND SUMMARY

1. 1 PURPOSE

The spacecraft reference trajectory defined in this report is de-

signed for the unmanned AS-206A Mission. The mission profile has

been designed such that the spacecraft mission objectives (Reference 1)

are satisfied without violating the spacecraft constraints or the mission

guidelines. The purpose of this report is to incorporate the mission

design refinements issued since the publication of the Apollo Mission

AS-206A Spacecraft Preliminary Reference Trajectory (Reference 2).

1. 2 SCOPE

The Apollo Mission AS-206A reference trajectory reflects the

mission refinements issued since the publication of Reference 2, in- "

corporates the known systems test requirements, and uses the current

guidance logic proposed for thrust vector and attitude control for the lunar

excursion module (LEM) powered flight maneuvers. The data presented

for the ascent-to-orbit phase are based on the AS-206A launch vehicle

trajectory listing transmitted to MSC by the Marshall Space Flight

Center (MSFC) through the Guidance and Performance Subpanel (Re-

ference 3).

This report consists of three volumes. Volume I presents the

mission requirements, a summary of the unclassified spacecraft input

data used in the generation of the mission profile, a description of the

major phases of the mission, and the trajectory analyses of pertinent

phases. Volume I also contains graphical and tabular time histories

of pertinent spacecraft attitude, position, motion, separation charac-

teristics, and MSFN coverage data.

Volume II of this report contains the classified trajectory listing

of the mission profile and the classified spacecraft input data specifi-

cations.

Volume III presents detailed time histories of radar range, range

rate, elevation, elevation rate, azimuth, azimuth rate, and four look-

angles for the MSFN ground stations available for operation during this

mission. . .

1

1.3 MISSION'^PROFIL'E SUMMARY

The AS-206A mission will be the first launch of a LEM spacecraft.

For mission simulation purposes, launch is assumed to occur at 13:00

hours Greenwich mean time (GMT), 1 April 1967, from launch complex

37B of the Kennedy Spaceflight Center.

Major mission events are illustrated in Figure 1-1. The mission

has been divided into 14 phases for discussion purposes:

1) Ascent-to-orbit .

2) S-IVB/space craft LEM adapter (SLA)/L,EM orbital coast

3) Spacecraft separation

4) Orbital cold-soak to first descent propulsion system(DPS) burn

5) First DPS burn

6) Orbital coast to second DPS burn

7) Second DPS burn/fire-in-the-hole (FITH) abort test/first ascent propulsion system (APS) burn

*8) Orbital coast to second APS burn ,

9) Second APS burn -

10) Orbital coast to third APS burn

11) Third APS burn

12) Orbital cold-soak to fourth APS burn • \

'13) Fourth APS burn

14) Final orbital coast

The Saturn IB launch phase includes the thrusting phases of the

S-IB and S-IVB stages. The boilerplate command and service module

(CSM) is jettisoned during S-IVB thrusting by the launch escape subsystem

(L.ES). S-IVB thrust termination occurs at a radius vector magnitude

of approximately 21, 440, 000 feet (perigee altitude of approximately 85

nautical miles) and a zero-degree inertial flight path angle. The velocity

at insertion is such that the radius vector magnitude at apogee is approxi-

mately 21, 650, 000 feet (altitude of approximately. 120 nautic'alT-'rniles). , ;.

These data correspond to the insertion state vector furnished to MSC by

MSFC through the Guidance and Performance Subpanel (Reference 3). .

The spacecraft is separated from the S-IVB/SLA combination during

the first revolution by the LEM/reaction control.system (RCS) thrusters..

MSFIST coverage of this mission event is provided by the Australia Apollo

tracking ship.

The first DPS burn is performed on the third revolution following

a spacecraft attitude-hold thermal soak (LEM X axis normal to the ecliptic

plane and Z axis toward the sun) of approximately 3 hours. The "Hohmann

descent" guidance philosophy is used for the first DPS burn. During this

burn (approximately 39 seconds), the AV available (approximately 150

feet per second) is used to raise the apogee altitude to approximately 180

nautical miles.

The second DPS burn/FITH abort test/first APS burn maneuver is

performed across the United States during the latter part of the third

revolution and the first part of the fourth revolution. The "lunar landing"

guidance philosophy is used for the second DPS burn (excluding the random

throttling phase). The major portion of the AV available (approximately

7000 feet per second) from this 757-second burn is dissipated out of the

orbit plane. The approximate apogee and perigee altitudes following the

random throttling sequence are 200 nautical miles and 155 nautical miles,

respectively.

FITH staging and the first APS burn are accomplished in a constant

inertia! attitude mode. The spacecraft attitude during the 5-second APS

burn dissipates most of the available AV (approximately 55 feet per second)

out of the orbit plane. The spacecraft orbit following this burn is essen-

tially unchanged from the orbit following the second DPS burn.

The second APS burn is performed over the United States near the

end of the fourth revolution. The "LEM ascent" guidance philosophy is

used for the second APS burn. Most of the available AV (approximately

5200 feet per second) resulting from this 395-second burn is dissipated

out of the orbit plane. At shutdown of the second APS burn, the space-

craft is on an orbit defined by an apogee altitude of approximately 203

nautical miles and a perigee altitude of approximately 163 nautical miles.

Following this mission event, and an attitude-hold orbital coast of approx-

imately 45 minutes, the third APS burn is performed while in sight of the

Carnarvon MSFN station. During the constant-attitude 5-second third

APS burn, the spacecraft thrust vector is oriented such that the orbit

perigee altitude is reduced to approximately 149 nautical miles. Approx-

imately 100 feet per second of AV is available from this burn.

The fourth APS burn is performed during the latter part of the

seventh revolution over the United States following an attitude-hold thermal

soak (LEM X axis normal to the ecliptic plane and Z axis toward the sun)

of approximately 3. 5 hours. During the fourth APS burn (approximately

8 seconds), the spacecraft attitude is maintained in an attitude-hold mode

by the abort guidance subsystem. The AV available from this burn

(approximately 160 feet per second) is used to decrease the spacecraft

perigee altitude to approximately 125 nautical miles.

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SPACECRAFT MISSION' REQUIREMENTS^"' -;

2. 1 SPACECRAFT TEST OBJECTIVES.

The purpose of Apollo mission AS-206A is.to conduct an unmanned

test of an operational L/EM for verification of subsystems operations. The

spacecraft test objectives are summarized in'this section.and presented in

Reference 1. ,

2. 1. 1 Primary Spacecraft Test Objectives . ; ':•

a) Verify operation of the following .L/EM subsystems:primary navigation and guidance control system(PNGCS), RCS, APS, and DPS.

b) Evaluate FITH abort test. • - . - • ' ' ""

2. 1. 2 .Secondary Spacecraft Test Objectives

a) Demonstrate APS operation at low propell ant quantities.

b) Demonstrate operation of the mission programmer.

2 . 2 MISSION PROFILE GUIDELINES - - ' . . . , '/

The mission profile .guidelines have been compiled from References

1, 4 through 8,, and supplemented by data furnished by the Manned Space-

craft Center (MSC) at technical coordination meetings.

2. 2. 1 Launch Vehicle Guidelines .

a) 'The nominal flight, azimuth shall be 72 degrees fromtrue North. • . ... '\ . .

7 •

b) The LES will be utilized to jettison the boilerplateCSM from the S-IVB/SLA/LEM combination whenthe dynamic pressure has decreased to less thanone pound per square foot.

c) The guidance command angular rate limitation shallbe one degree per second for both pitch and yawsteering.

2. 2. 2 MSFN Coverage/Flight Control Guidelines ;

a) Continuous MSFN tracking, telemetry, and groundcommand backup coverage are required from liftoffto S-IVB thrust termination (orbital insertion) plus3 minutes.

b) Continuous telemetry, tracking, and ground command; backup coverage are required for all LEM descent and

ascent stage firings and pre-ignition attitude maneuvers.This coverage should commence at least 2 minutesprior to the programmed ignition and should not endprior to confirmation of shutdown.

c) Telemetry and tracking coverage are required at leastonce per revolution during orbital coasts and thermalsoaks.

d) Telemetry, tracking, and ground command backupcoverage are required at least 3 minutes prior tospacecraft separation. This coverage should continueuntil the LEM can radiate to a MSFN station.

i e) State vector update capability is required near theCarnarvon MSFN station.

f) For mission events requiring command coverage bymore than one MSFN station, it is desired that thecoverage overlap be at least 30 seconds.

g) The FITH abort test/first APS burn shall be performedat an orbital position so that at least three MSFN sta-tions with data record capability can receive data fromthis demonstration.

h) The LEM very high frequency (VHF) slant rangeacquisition limit is approximately 740.nautical miles.

2. 2. 3 LEM Attitude Guidelines

a) Shortly after orbital insertion, the S-IVB/SLA/LEMcombination shall be maneuvered such that the X axislies in the plane of the local horizontal and the minusZ axis is directed downward along the local vertical.This attitude shall be maintained in an orbital rate modeuntil the start of the separation sequence to enhancecommunications.

b) A ±0. 3-degree deadband attitude-hold mode shall bemaintained during the separation sequence.

t c) Shortly after separation from the SLA. the LEM shallbe oriented such that the X axis is aligned to within±15 degrees of a normal to the ecliptic plane and theplus Z axis is directed toward the sun. This attitudeshall be maintained for approximately 3 hours forthermal soak of the DPS injector face.

d) During coast periods when the LEM attitude is nototherwise constrained, the minus Z axis shall be

10

directed downward along the local radius vectorto enhance communications and to provide backupfor VHF pulse code modulation (PCM) telemetry.

e) All orbital coasts requiring specific orientationsshall be maintained in the ±5-degree deadband.attitude-hold mode.

f) . Shortly after the third APS burn, the LEM shall beoriented such that the X axis is aligned to within ±15degrees of a normal to the ecliptic plane and the Zaxis is directed toward the sun for thermal soak ofthe APS injector face and RCS quad III. This attitudeshall be maintained for approximately 3 hours.

2. 2. 4 LEM Propulsion Guidelines

a) LEM separation from the S-IVB/SLA combinationshall be accomplished during the first revolutionusing four RCS thrusters firing for 18 seconds toprovide plus X axis translation. The LEM/SLArestraining, straps shall be severed 2 seconds afterRCS ignition. This RCS firing time provides adequateclearance between the stowed landing gear and theSLA petals should one. of the RCS engines fail to fire.

b) Ullage settling maneuvers shall precede all DPSfirings using four RCS thrusters to provide plus Xaxis translation. These ullage maneuvers shall be6 seconds in duration. . .

.• - _ . • . . •- - - f \- • • -

c) Ullage settling maneuvers shall precede the .second,third, and fourth APS firings using four RCS thrustersto. provide plus X axis translation. These ullagesettling maneuvers .shall be 2. 5 seconds in duration.

d) ' The start sequence of each DPS burn shall be asfollows: ignition and the first 3 seconds of the burnshall be at the 30-percent thrust setting; the next23 seconds shall be at the 10-percent thrust setting.

e) The first DPS burn profile shall consist of the startsequence described in the above paragraph followedby approximately 7 seconds of thrusting at the 92. 5-percent thrust setting. Shutdown of the first DPSburn shall occur approximately 33 seconds afterignition.

f) The second DPS burn profile shall consist of thestart sequence previously described followed by379 seconds at the 92. 5-percent thrust setting; forthe next 300 seconds, the thrust shall decrease.linearly from the 60- to the 10-percent thrust

11

setting. The random throttling portion of the burnshall begin following the thrust decay to the 10-percentthrust setting and shall be 52 seconds in duration. Therandom throttling profile shall consist of 10-secondintervals at the 10-, 50-, 30-, 40-, and 20-percentthrust settings followed by a 2-second burn at the 92. 5-percent thrust setting.

g) The FITH abort test is to be initiated immediately afterthe 92. 5-percent thrust setting of the random throttlingphase.

•h) The first APS burn (initiated with the FITH abort test)shall be approximately 5. seconds in duration.

i) The second APS burn shall be approximately 435seconds induration.

j) The third APS burn shall occur approximately 45minutes after shutdown of the second APS burn.

k) The third and fourth APS burns shall be approximately5 seconds in duration.

1) Demonstration of RCS operation with propellant trans-ferred from the APS tanks shall begin 2 seconds afterthe start of the second APS burn and will continue untilshutdown.

m) The fourth APS burn duration shall deplete the APSpropellants to less than 5 percent of the total available.

2. 2. 5 General Mission Guidelines

a) The predicted orbital lifetime for the spent descentand as.cent" stages should not exceed 3 months.

b) LEM separation from the S-IVB/SL.A shall occurduring the first revolution in the vicinity of theCarnarvon MSFN" station.

c) LEM landing gear deployment, shall occur followingthe orbital cold-soak prior to the first DPS burnwhile the spacecraft is in contact with a MSFN station.

d) Pre-ignition attitude orientation maneuvers shalloccur within MSFN coverage by a station withcommand backup capability.

e) The fourth APS burn will be guided by the abortguidance subsystem (ACS) in the attitude-holdmode.

12

... ,. 3. SUMMARY OF : INPUT DATA . - - : r . ., .

The input data contained in this section include the pertinent MSFN

station specifications and the unclassified quantitative description of the

spacecraft. The spacecraft specifications classified as confidential can

be found in Volume II of this report. These data form the basis for the .

spacecraft reference trajectory in support of Apollo Mission AS-206A.

3. 1 -' SATURN IB LAUNCH VEHICLE "

The Saturn IB launch vehicle which will be, used to insert the S-IVB/

SLA/LEM combination into the initial orbit is comprised of the S-IB and

S-IVB stages. The official launch vehicle data and launch vehicle refer-

ence trajectory. (Reference 9) will be furnished by Marshall Space Flight

Center '(MSFG). The Saturn IB launch vehicle is illustrated in Figure 3-1.

3.2 . SPACECRAFT (LEM-1) . ., .

The spacecraft weight statement was obtained from Reference 8 and

is presented in Volume II of this report. The propulsion characteristics

for both the ascent, and descent stages of the. spacecraft are presented in

Volume II of this report. The spacecraft reference dimensions are pre-

sented in Figure 3-2. The LEM-i outboard profile ,is illustrated in

.Figure 3-3. • • - . - • . .

3 . 3 MSFN STATIONS . . . . , .

The locations and capabilities of the-MSFN stations that are planned

to be available for support of Apollo Mission AS-206A were obtained from

References 5 and 10, respectively. These data are summarized in Table

3-1. The station coordinates given are based on the Fischer ellipsoid

model described by an equatorial earth radius of 6, 378, 166. 0 meters, a

polar earth radius of 6, 356,784. 284 meters, and an earth flattening ratio

of 1/298. 3. MSFN station altitudes are referenced to this ellipsoid and

include any known geoidal separation.

The criteria for selecting the locations of the three Apollo tracking

ships which are listed in Table 3-1 are discussed in the following para-

graphs. The Australia Apollo tracking ship (Ship Number One) is planned

for location off the western coast of Australia. The coverage afforded by

13 .

the Carnarvon tracking station and this tracking ship will be used for the

spacecraft separation sequence, state vector updates, the first DPS burn,

:and the third APS burn. The location of this ship is such that the 30-sec-

ond coverage overlap requirement between this station and Carnarvon and

the 3 minutes of coverage prior to separation requirement are satisfied.

This location also provides --at-least 2 minutes of coverage prior to the

first DPS burn and the third APS burn. Summaries of the study data used

in the selection of the location of this ship are illustrated in Figures 3-4

and 3-5.

The California Apollo tracking ship (Ship Number Two) is to be

located off the western coast of the continental United States. Coverage

provided by this station is required for the second DPS burn and the sec-

ond and fourth APS burns. Because of the large central angle traversed .

by the spacecraft during the second DPS burn, it was required that the

location of this ship be positioned to increase coverage for this mission

event, yet provide the necessary coverage for the second and fourth APS

burns. To satisfy these requirements, this station was located as far off

the western coast of the United States as possible, while still retaining a

30-second coverage overlap with the Texas MSFN station.

The Bermuda Apollo tracking ship (Ship Number Three) is located

so that it provides coverage from 30 seconds prior to S-IVB shutdown

(orbital insertion) to at least 3 minutes after S-IVB shutdown. A summary

of the study.data used in the selection of the position of this ship is pre-

sented in Figure 3-6. The data presented in Figure 3-6 may. also be used

to adjust the position of this ship and to determine the location of any

additional coverage sources which may be required for support of possible

contingency situations.

3. 4 EARTH CONSTANTS AND CONVERSION FACTORS

The earth constants and conversion factors presented on the follow-

ing page were extracted from Reference 4 and have been used in the

generation of the spacecraft reference trajectory.

14

3. 4. 1 Earth .Constants

Rotational rate

Equatorial radius

Average radius

Gravitational parameter (|a )

4. 37526902 x 10"3 rad/min

0. 417807416 x 10~2 deg/sec .

0.729211504 x 10"4 rad/sec

2. 092573819 x 10? ff' . :

2. 0909841 x 107 ft

5. 53039344 x 10"3 er3/min2

11. 46782384 x 103 er,3/day2

3. 986032 x 105km3 /sec2-

1.407653916 x 1016 ft3 /sec2

Coefficients of potential harmonics

J term .(second)

H term (third) ' '

D term (fourth)

Earth flattening

3. 4. 2 Conversion Factors

Kilometers per foot

Kilometers per nautical mile

Feet per nautical mile

Weight-to-mass ratio

Mass-to-weight ratio

Feet per earth equatorialradius

Nautical miles per earthequatorial radius

-3-1.62345 x 10 nd

-0.575 x 10~5 nd

0. 7875 x.10"5 nd

1/298: 3;nd

0.'3048 x 10~3km/ft

. ,1..852 krn/n mi

.6076.115486 -ft/n mi

32.'17404856 Ib/slug

0.031080950 slug/lb

2. 092573819 x 10? f t /er

3443. 93358 n mi/er

15

3. 5 SPACECRAFT AND ATTITUDE REFERENCE COORDINATE '-.SYSTEMS

The spacecraft and attitude reference coordinate systems are illus-

trated in Figures 3-7 through 3-9. The spacecraft attitude is measured

by the pitch, yaw, and roll angles required to rotate from the reference

..system .to the. current spacecraft orientation. . , . . ______

3. 5. 1 Spacecraft Coordinate System, X „, Yg, Z_

The LEM coordinate system (see Figure 3-7) is an orthogonal,

right-handed system coincident with the spacecraft axes. The Xg axis

extends through the upper docking tunnel. The Zg axis extends along the

crew line -of -sight and the Yg axis completes the right-handed system.

3. 5. 2 Launch Site Inertia! Reference System, XT, YT, Z

This orthogonal, right-handed coordinate reference system (see

Figure 3-8) coincides with the launch site at the time of launch. The Xy

axis extends downrange in the direction of the launch azimuth. The Z

axis extends upward along the astronomical vertical and the YT axis

completes the right-handed system.

3. 5. 3 Launch Site Rotating Reference System, XR, Y ', ZR

This coordinate reference system (see Figure 3-8), which rotates

with the earth,is coincident with the XT, YT, Z system at launch.

3. 5. 4 Relative Vehicle Coordinate System, XR-,, YRV>

This coordinate system (see Figure 3-9) is an orthogonal, right-

handed system centered at the primary vehicle. The X-...... axis extends_ • K. V

in the direction of motion. The YRV axis extends upward along the prima-

ry vehicle position vector. The Z_v axis completes the right-handed

system.

16

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Figure 3-1. Saturn IB Launch Vehicle Outboard Profile

18

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19

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Figure 3-3. LEM-1 Outboard Profile

20

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Figure 3-4. Australia Apollo Tracking Ship Placement Study/Spacecraft Separation

TOTAL COVERAGEOF SELECTED SHIPPOSITION

ILATITUDE OF AUSTRALIAAPOLLO TRACKING SHIP(DEC SOUTH)

SELECT D SHIP POSI ION

MINIMUM OVERLAP TIME (30 SECONDS)

92 14 96 98 100 103 104 106

AUSTRALIA SHIP LONGITUDE (DEC EAST)

Figure 3-5. Australia Apollo Tracking Ship Placement Study/First DPS Burn

21

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24

4. MISSION DESCRIPTION

The spacecraft reference trajectory for Apollo Mission AS-206A ,

has been designed to satisfy the mission test objectives, (Reference 1) .

and known test requirements without-violating the mission profile .guide-i

lines (Section 2. 2). This mission profile- uses the current L/EM guidance

philosophy for thrust vector and attitude control during the spacecraft

first DPS, second DPS.and second APS burns. The spacecraft first,

third, and fourth APS burns are accomplished in an attitude-hold mode.

The description of the AS-206A mission simulation and the results of

pertinent mission planning studies and mission analyses are presented

in this section.

4 . 1 ASCENT-TO-ORBIT : . . . ; . '

Launch of Apollo Mission AS-206A is planned to occur from launch

complex 37B of the Kennedy Spaceflight Center. ; F'br mission simulation

purposes, launch is assumed to occur at 13:00 hours GMT (08:00 hours

EST) on 1 April 1967. " . - •• '

The ascent-to-orbit profile presented in this report is based upon

an MSEC trajectory listing transmitted to MSC through the Guidance and.

Performance Subpanel (Reference 3). The .ascent-to-orbit phase of the

mission profile utilizes a 72-degree flight azimuth from true North and

includes the burn of the S-IB stage and the burn of the S-IVB stage. The

boilerplate C.SM is jettisoned when the dynamic pressure decreases to

less than one pound per square foot. . " . . ' • :

Insertion into an elliptical earth orbit occurs 603.44 seconds after

liftoff. The state vector at insertion is described as follows:

• Inertial velocity magnitude of 25, 658 feet per second

• Radius vector magnitude of 2.1, 440, 000 feet(approximately 85 nautical miles)

Inertial flight path angle of 0. 0 degree

Orbital plane inclination-of 31. 6.1 degrees. -

25

The ascent-to-orbit phase presented in this report has been de-

signed to satisfy the previously mentioned conditions. MSFN coverage is

provided by the Cape Kennedy, Grand Bahama, Grand Turk, Bermuda, and

the Bermuda Apollo tracking ship stations. It should be noted that the data

for this mission phase are presented for completeness only; the official

launch vehicle^ data will be published by MSFC in the launch vehicle ref-

erence trajectory (Reference 9).

4. 2 S-IVB/SLA/LEM ORBITAL COAST

Shortly after insertion into the initial orbit, the S-IVB/SLA/LEM

combination is maneuvered so that the spacecraft X axis lies in the plane

of the local horizontal and the minus Z axis is directed along the radius,

vector toward the earth. This attitude is maintained by the S-IVB in an

orbital rate mode until 120 seconds prior to LEM separation to enhance

LEM/MSFN communications. SLA petal deployment occurs during the

first revolution (30 minutes after .orbital insertion).

4.3 SPACECRAFT SEPARATION

LEM separation from the S-IVB/SLA combination occurs near

apogee during the first revolution, approximately 10 minutes after SLA

petal deployment. The selection of the in-orbit position for this mission

event is such that continuous MSFN ground command coverage is provided

by the Australia Apollo tracking ship and the Carnarvon station. Sixty

seconds after acquisition the S-IVB begins an attitude-hold maneuver,

maintaining a constant inertial attitude through the spacecraft separation

sequence. Two minutes after the beginning of the inertial attitude-hold

period, four LEM/RCS thrusters are ignited to provide plus X axis

translation. At RCS ignition, the spacecraft plus X axis is in the orbital

plane approximately 9 degrees above the local horizontal. Two seconds

after RCS ignition, the SLA/LEM restraining straps are severed and the

LEM begins the withdrawal from the SLA, Eighteen seconds after ignition

RCS thrusting is terminated. The separation characteristics (LEM-to-

S-IVB relative velocity and distance) are presented in Figures 4-1 and

4-2. The total coverage provided by the two MSFN stations for this event

is approximately 8 minutes and 40 seconds.

26

4. 4 ORBITAL COLD-SOAK TO FIRST DPS BURN " ' ' : :'

Ten seconds after shutdown of the four RCS engines in the. space-

craft separation sequence, the LEM is commanded to perform an attitude

maneuver to align the X axis normal to the ecliptic plane and to direct

the Z axis toward the sun. The spacecraft maintains this inertial attitude

in the ±5-degree deadband mode for the duration of the cold-soak (ap-

proximately 3 hours). The spacecraft attitude orientation during the

cold-soak period is illustrated in Figure 4-3.

Since the sun's position relative to the spacecraft is dependent on

the launch date and time, the maneuver included in this mission profile is

valid only for the assumed launch date and time (13:00 hours GMT,

1 April 1967). The spacecraft inertial attitude angles which satisfy this

thermal cold-soak requirement are presented in-Figure 4-4 as a function

of launch date and time for the second quarter of 1967.

4. 5 FIRST DPS BURN

Upon acquisition of tracking by the Australian tracking ship on

the third revolution, an'orientation maneuver is initiated to align the

spacecraft axes to the desired orientation for'the first DPS burn (see

Figure 5-14). This maneuver is executed in the 5-degree per second

rate mode. Shortly after the termination of this maneuver, the LEM

landing gear is deployed. Approximately one minute has been provided

prior to RCS ignition for MSFN verification of gear deployment.

The first DPS ignition is preceded by a 6-second RCS plus X axis

translation for ullage settling. The first DPS burn start sequence con-

sists of ignition and the first 3 seconds of burn time at the 30-percent

thrust setting, followed by 23 seconds at the 10-percent thrust setting.

This start sequence is followed by a one-second buildup to, and approxi-

mately 11. 5 seconds at, the 9.2. 5-percent thrust setting (see Figure 4-6).

The "Hohmann descent" LEM guidance philosophy is used to con-"

trol the spacecraft attitude and attitude rates during this burn (Reference

11). This guidance philosophy is characterized by burnout at a zero-

degree flight path angle. : •

27

At shutdown of the first DPS burn, the spacecraft is on'an orbit

defined by a predicted apogee altitude of 180 nautical miles, a predicted

perigee altitude of 119 nautical miles, and an orbital period of 90 minutes.

The orbit is inclined 31. 62 degrees to the equator and the longitude at* •' . " ~ "

burnout (perigee) is 96. 5 degrees east. The continuous MSFN ground

command coverage of the first DPS burn is provided by the Australia

tracking ship and the Carnarvon station. The total coverage time is

approximately 9 minutes and 30 seconds.

* It was determined that the second DPS burn would be the least

sensitive to attitude errors if the line of apsides of the orbit following

the first DPS burn would bisect the central angle traversed by the space-

craft during the second DPS burn. An orbit defined by a 120-nautical

mile perigee and a 180-nautical mile apogee resulted in a first DPS

burn duration which satisfied the mission guidelines and provided con-

tinuous MSFN command coverage during the first and second DPS burns.

4. 6 ORBITAL COAST TO SECOND DPS BURN

Following shutdown of the first DPS burn, the spacecraft begins

an orbital coast of approximately 35 minutes. Shortly after the begin-

ning of the coast, a LEM maneuver to orient the X axis in the plane.ofi •

the local horizontal and the minus Z axis toward the geocenter is per- .

formed at the 5-degree per second attitude-rate mode. The spacecraft

is oriented in this manner to enhance LEM/MSFN communications.

This attitude is maintained in the ±5-degree attitude-hold mode.

4.7 SECOND DPS BURN/FITH ABORT TEST/FIRST APS BURN

4^7. 1 Second DPS Burn

At a mission time of approximately .4 hours and 33 minutes the LEM

begins an orientation maneuver to the desired inertia! attitude (Figure

5,-18) for the second DPS burn. The attitude maneuver is performed in

the 5-degree per second rate mode. Approximately 70 seconds after the

completion of the attitude orientation maneuver, a 6-second RCS maneuver

is initiated to provide plus X axis translation for ullage settling immedi-

ately prior to DPS ignition. The desired thrust profile, for the second

DPS burn (Figure 4-7), consists of the start sequence described in

28

Section 4.'5, followed by a 379-second burn" phase- at the -92. 5-rpercent'

thrust setting and a 300-second linear decay from 'the 60- to the 10- •-" '

percent thrust setting. Upon completion of-the linear decay, a 52-second

random throttling sequence is initiated. This throttling profile consists '

of ten-second intervals at the 10-, 50-, 30-, 40-, and'20-percent thrust

settings, respectively. The random throttling profile ends with a two- •

second burn at the 92. 5-percent thrust setting; This burn profile is

typical of the lunar landing mission maneuvers that may be performed

from the completion of lunar de-orbit to lunar;touchdown.,., , .

The guidance philosophy used to provide control of the second DIPS '

burn thrust vector magnitude and direction utilizes multiple targeting •'

conditions (Reference 11). .Each targeting condition consists of the de?-

sired position, velocity, and acceleration vectors and the burn duration,

for the spacecraft to satisfy these conditions. It should be rioted that

the random throttling sequence, FITH staging, arid the first APS.-burn :.

are performed in an attitude-hold mode. , . . . - , .

The second DPS thrust profile resulting from the guidance com-

mands is compared to the desired profile (from the mission guidelines) ,

in Figure 4-7. • : . . , . . . - ' ; .

At shutdown of the second DPS burn, the LEM is- on an orbit with

a predicted perigee altitude of 155 nautical miles, a predicted apogee

altitude of 200 nautical miles, 'a period of approximately 90. 9 minutes,

and an orbital inclination of approximately 30. 5 degrees. An open loop :

analysis to determine the effects of an assumed IMU drift (Reference 7) on

various spacecraft orbit dimensions following the second DPS burn was

performed. The data, generated from this study-are presented in Figure

4-8 and were used in the selection bf'the desired o.rbit dimensions fol-

lowing, the second DPS burn.

The majority of the AV available from this burn (approximately

7, 000 feet per second) is dissipated out of the orbit plane. Approximately-

one-percent of the useable DPS propel-lant remain after the second DPS

burn. - - . , - . . •

29

MSFN coverage of the second DPS burn is provided by the California ;

Apollo tracking ship, Goldstone, Texas, Cape Kennedy, Guaymas, White ;

Sands, Grand Bahama, Eglin, and Bermuda stations. : . :. ,

4. 7. 2 FITH Abort Test/First APS Burn

The FITH a'bort test is initiated immediately at the completion of

the random throttling phase of the second DPS burn with simultaneous

LEM staging and APS ignition signals. The duration of the first APS

burn is 5. 24 seconds. The in-orbit position of this event has been selec-

ted such that MSFN ground command coverage and data record capability

are furnished by the Cape Kennedy, Grand Turk, Bermuda, Grand Bahama,

and Eglin stations. In the mission simulation, 1, 180 pound-seconds

of impulse were delivered to both the LEM ascent and descent stages at

the APS ignition to simulate APS plume impingement effects. Separation

characteristics for the ascent and spent descent stages are presented in

Figures 4-9 and 4-10. At FITH staging, the spacecraft X axis is approxi-

mately normal to the orbital plane. At shutdown, the ascent stage is oil

an orbit defined by a predicted perigee altitude of 152. 5 nautical miles,

a predicted apogee altitude of 200. 0 nautical miles, a period of approxi-

mately 90. 9 minutes, and an inclination of approximately 30. 5 degrees.

The spent descent stage is left in an orbit with a predicted apogee altitude

of 200. 2 nautical miles, a predicted perigee altitude of 153. 4 nautical

miles, an inclination of approximately 3.0. 5 degrees, and a period of

90. 9 minutes. The maximum estimate of the lifetime of the spent descent

stage is 49 days (see. Table 4-2). _—' '

4. 8 ORBITAL COAST TO SECOND APS BURN

Shortly after the first APS burn, a LEM ascent stage attitude ma-

neuver is initiated to align the X axis in the plane of the local horizontal

and to direct the minus Z axis toward the geocenter. This maneuver is

accomplished in the 5-degree per second maneuver mode. This orienta-

tion is maintained in the ±5-degree attitude-hold mode throughout this

coast (approximately 1 hour and 20 minutes).

4. 9 SECOND APS BURN

Following acquisition by the California tracking ship during the

30

four th revolution, an attitude-maneuver is initiated to orient the LEM'

ascent stage to the desired ignition attitude (see Figure 5-22). This

maneuver is performed in the 5-de'gree pe:r second maneuver mode.' •

Two minutes after the acquisition of tracking, the RCS thrusters are

ignited for 2. 5 .seconds of plus X. axis translation for ullage settling.

Immediately after the 2. 5-second RCS burn, the second APS burn is

ini t iated.

The "LEM ascent" guidance, philosophy is used to control the. ;.

spacecraf t attitude, attitude rates., and. second APS burn duration. „ This -

philosophy is characterized by the use of a desired velocity vector at

burnout and 'an intercept time. The burn duration is varied so that the .

velocity vector requirements, are satisfied at the intercept time. The

philosophy used to target this burn consisted of varying the intercept

parameters such that the desired burn time (approximately 395 seconds) ,

and orbital .conditions were met. .It was required that the second APS-

burn duration be reduced from that specified in Reference 1 (approxi-

mately 435 seconds) in order that sufficient propellants remain for .the.

required third and, fourth APS burns. . . . ; .

At shutdown of this burn, the spacecraft orbit is defined by a pre-

dicted apogee altitude of 203. 1 nautical miles, 'a predicted perigee alti-

t u d e of 163. 3 nautical miles, and a period of approximately 91. 2 minutes.

The orbital plane-is inclined approximately 31.4 degrees to the equator.

The desired orbit dimensions at second APS burn shutdown were chosen

so that the orbital lifetime of the ascent stage could be significantly

reduced during the third and fourth APS burns and to reduce the effects

of the assumed IMU drift (Figure 4-10). Continuous MSFN coverage of

this burn is furnished by the California Apollo tracking ship , the Texas,

White Sands, Guaymas, Goldstone, and Point Arguello stations. The

total coverage time is approximately 13 minutes.

Two s-econds after initiationyof the second APS-burn, a test of RCS

operation is initiated. This test is conducted to determine the operation

of the RCS using propellants fed from the APS propellant tanks. The

RCS is used for stabilization of the ascent stage during APS operation.

This test is continued until shutdown of the second APS burn.

31

As in the second DPS burn, the majority of the AV available during

this burn (approximately 5, 200 feet per second) is.dissipated out of the

orbit plane to maintain orbit dimension control.

4. 10 ORBITAL, COAST TO THIRD APS BURN

Shortly after shutdown of the second APS burn, the LEM is com-

manded to maneuver to the inertial attitude described in Section 4. 8 at

the 5-degree per second attitude maneuver mode. This attitude is main-

tained in the ±5-degree deadband mode until acquisition of tracking by

the Australia tracking ship during the fifth revolution (approximately

53 minutes after shutdown of the second APS burn).

4. 11 THIRD APS BURN

Upon acquisition of tracking by the Australia Apollo tracking ship,

the LEM is commanded to maneuver to the desired ignition attitude for

the third APS burn in the 5-degree per second attitude orientation mode.

Two minutes after acquisition of tracking, a 2. 5-second RCS maneuver

is performed to provide plus X axis translation for ullage settling. Im-

mediately after the ullage maneuver, the third APS burn (5. 24 seconds

in duration) is initiated. A constant inertial attitude was selected for this

burn to decrease perigee altitude (thus orbital lifetime). This burn

results in an orbit with an apogee altitude of approximately 220. 1 nautical

miles, a perigee altitude of approximately 149. 4 nautical miles, a 91. 2-

minute period, and an orbit plane inclination of approximately 3 1. 4 de-

grees. Approximately 10 minutes and 30 seconds of continuous MSFN

coverage is furnished for this burn by the Carnarvon and Australia

tracking ship stations.

4. 12 ORBITAL COLD-SOAK TO FOURTH APS BURN

Shortly after completion of the third APS burn the LEM is maneu-

vered such that the X axis is normal to the ecliptic plane and the Z axis

is directed toward the sun for cold-soak of the APS injector face. This

maneuver is accomplished in the 5-degree per second attitude orientation

mode. This attitude is maintained in the ±5-degree deadband mode for

approximately 3 hours and 30 minutes.

32

4. 13 FOURTH APS BURN ." V

Following acquisition by the California tracking ship during the

seventh revolution, an orientation maneuver is initiated in the 5-degree

per second mode to align the spacecraft axes to the desired ignition

attitude for the fourth APS burn. Two minutes after acquisition, 2. 5

seconds of RCS thrusting is initiated to provide plus X axis translation

for ullage settling. This is followed immediately by the fourth APS burn

of approximately 8.2 seconds which depletes all of the available APS

propellarits with the exception of the one percent flight performance . ;

reserves. • • ' . . .

The inertial attitude held during the APS burn further decreases

the perigee attitude. At APS shutdown, the spent spacecraft is on an

orbit characterized by a predicted perigee altitude of 124. 8 nautical

miles, a predicted apogee altitude of 199. 0 nautical miles, an orbital

period of approximately 90. 4 minutes, and an orbital inclination of

approximately 31. 4 degrees. The estimated lifetime of the spacecraft

in this orbit is approximately 19 days.

4. 14 FINAL ORBITAL COAST : . • .

The majority of the mission objectives are completed at this phase

of the mission; however, additional spacecraft tests may be scheduled

after this time. An additional 3 hours have been added after the com-

pletion of the fourth APS burn to provide data for planning any additional

spacecraft testing.

4. 15 SPACECRAFT ORBITAL LIFETIME ESTIMATES

The estimated orbital lifetimes of the various LEM-1 configurations

are presented in Table 4-2. The configurations which have been analy-

zed are defined below:

Configuration 1

The LEM-1 after separation from the S-IVB/SLA combinationand prior to propulsion tests.

33

Configuration 2

The-spent descent stage after the FITH abort test.

Configuration 3

The spent ascent stage after shutdown of the fourth APS burn.

•. Ballisticrcoefficients, W/C.-.A (weight divided by.the orbital drag

coefficient and the frontal area), were calculated for each of the configu-

rations. An orbital drag coefficient of 2. Owas assumed. Various con-

figuration orientations were analyzed to determine the maximum and .

minimum frontal areas which each configuration could exhibit normal to

the velocity vector. Table 4-1 presents the results of this analysis.

Table 4-1. Ballistic Coefficients

Configuration

12233

Area (ft2)

* 200(Maximum) 200(Minimum) 80(Maximum) 190(Minimum) 125

Weight (Ib) '

33, 4755, 1955, 1956, 0576, 057

W/C-A ( lb / f t 2 )JJ83. 69'12. 9932. 4715.9424. 23

*The frontal area of the LEM does not change appreciably.

These ballistic coefficients, the applicable orbital characteristics,

and information presented in Reference 10 were used to calculate the

orbital lifetime estimates presented in Table 4-2.

Table 4-2. Spacecraft Orbital Lifetime Estimates

Orbital Lifetime (days)C onfigur ation Minimum Maximum

1 4 42 19 493 13 19

34

50

40

u*.UJ

•g 30

t—

—OUJ

S 20

UJ

10

>-I—

o

SPACECRAFT SEPARATION OCCURS52 MIN 5 SECFROM LIFTOFF

-4 0 4 . 8 12 16

TIME FROM SEPARATION (SEC)

•20

Figure 4-1. Spacecraft Separation; Relative Velocity and DistanceDuring RCS Thrusting

RCS SHUTDOWNOCCURS 52 MIN 21 SEC •FROM LIFTOFF

100,000

80,000

Z 60,0002£0

S 40,000

20,000

0 L.

12.0

100

> 80tb

- y 60

40

20

•RCS SHUTDOWN

RELATIVEVELOCITY

0 U»I

///

^.RELATIVE./DISTANCE

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

TIME FROM RCS SHUTDOWN(HR:MIN)

Figure 4-2. Spacecraft Separation; Relative Velocity and DistanceHistory for Approximately 100 Minutes

35

. 240

15APR .1MAY / • I5MAY . '.' IJUN

• • ..'; ' '. ' • LAUNCH DATE (1967) "

15JUN I JUL

1APR 15APR 1MAY 15MAY IJUN

LAUNCH DATE (1967)

15JUN I JUL

J

OS

Ou

GREENWICH MEAN TIME AT LAUNCH (HRS)

IAP8 I5APR IMAY I5MAY IJUN

LAUNCH DATE (1967)

I5JUN I JUL

Figure 4-4. Cold -Soak. Attitude Orientation Historyfor Launches, Second Quarter, 1967

37

100

80

60

40

20

\DESIRED THRUSTGUIDANCE COMMANDED THRUST \

0 . 4 6 8 10

TIME FROM DPS IGNITION (MIN)

12 14

Figure 4-5. First DPS Burn Thrust Profile

100

80

1 60

40

20

10 20 . 30 40 50

TIME FROM DPS IGNITION (SEC)

Figure 4-6. Second DPS Burn Thrust Profile

38

o .-

If 300

170

OZ '-: ; •>=•' " 160

120

80

180 190 230,200 . 210 220

DESIRED APOGEE FOLLOWING SECOND DPS BURN ( N Ml)

250

NOTE:1 . DESIRED STUDY PERIGEE IS 175 N Ml2. ASSUMED IMU DRIFT OF .1 .2 DEC APPLIED

IN DIRECTION OF MAXIMUM ORBIT SENSITIVITY.

Figure 4-7. Effect of Assumed IMU Drift on Orbit Dimensionsfollowing Second DPS Burn

FITH STAGING OCCURS4 HRS 47 MIN 56 SECFROM LIFTOFF

. 200

t '60LU

u

1/1 120

LU

Lu 80a:

• 40

0

i i

i T

r

*TIV

E V

ELO

CIT

Y (

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ME IMPIN

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X'

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5 /APS IG

CEMENT E

' DISTAN

^ - - - - - -

^PS SHUTD

NITION-

FFECTS ,

^:E

OWN-*-

/

f

^

2 ' 3 4 5

TIME FROM FITH STAGING (SEC)

Figure 4-8. FITH Staging; Relative Velocity and Distance HistoryDuring first APS Burn"

39

AN

CE (F

T)

7000

IV

LAT

IVE

VE

LOC

ITY

(F

FITH STAGING OCCURS4 MRS 47 MIN 56 SECFROM LIFTOFF

20 40 60 80

TIME FROM FITH STAGING (SEC)

100

Figure 4-9. FITH Staging; Relative Velocity and DistanceHistory to 100 Seconds

170 10 190 200 2JO 220 230 240DESIRED APOGEE FOLLOWING SECOND APS BURN ( N Ml)

2 0

NOTE:DESIRED STUDY PERIGEE IS 175 N Ml

2. ASSUMED IMU DRIFT OF 1.7 DEGREES APPLIEDIN DIRECTION OF MAXIMUM ORBIT SENSITIVITY

Figure 4-10. Effect of Assumed IMU Drift on Orbit DimensionsFollowing Second APS Burn

40

5. NOMINAL"TRAJECTORY DATA

, This section contains pertinent trajectory parameter histories of the

nominal mission profile. These data, in tabular and graphical form, are

based upon the trajectory data presented in Volume II'of this report.

Volume II also contains definitions of the trajectory parameters presented

i n this section. • ' • ' . , -

5. 1 MISSION PROFILE DATA ; :

The time sequence of events for Apollo Mission AS-206A is presented

in Table 5-1. Figures 5-1 through 5-4 present the earth ground track for

the entire mission. Orbital characteristics of the spacecraft coast phases

are presented in Table 5-2. Earth shadow data (daylight-darkness) are

illustrated in Figures 5-1 through 5-4 and tabulated in Table 5-3.

5.2 .TRAJECTORY PHASE DATA

Discrete events summaries and pertinent time history ;data of the

spacecraft position, motion, • and attitude are presented for the major ' '

mission phases as follows:

Mission Phase Table • • Figures

Ascent-to-Orbit 5-4 5-5 through 5-8

Spacecraft Separation 5-5 5-9 through 5-11

First DPS Burn.. ' . 5-6 5-12 through 5-15 :

Second D P S Burn/FITH'Staging/ . . . . . . .First APS Burn . 5.-7 5-16 through 5-19

Second APS Burn 5-8 . . . . 5-20 .through 5-23

Third APS Burn 5-9 5-24 through 5-26

Fourth APS Burn 5-10 . 5-27 through 5-29

The spacecraft attitude angles presented in the figures are refer-

enced to a launch-centered inertial-coordinate system. This-coordinate

system and the spacecraft axis system are illustrated in .Figures 3-7 and

3-8.

41

Table 5-1. Time Sequence of Events

• ' . - - . - . - . . - . - . - Time From LiftoffEvent (hr:min:sec)

Ascenti-to-Orbit

-- Liftoff - - . -. . 00:00:00.0S-IB Inboard Engines Shutdown 00:02:17.6S-IB Outboard Engines Shutdown 00:02:23. 6S-IB Jettison/S-IVB Ignition 00:02:29. 1Jettison Boilerplate CSM 00:02:49. 1S-IVB Shutdown into Elliptical Earth Orbit 00:10:03. 4

S-IVB/Spacecraft Orbital Coast

S-IVB Shutdown into Elliptical Earth Orbit 00:10:03. 4SLA Petal Deployment 00:40:03. 4

Spacecraft Separation

RCS Ignition 00:52:02. 8Sever Restraining Straps 00:52:04.8RCS Shutdown 00:52:20. 8

Spacecraft Orbital Cold-Soak

RCS Shutdown 00:52:20. 8Begin Cold-Soak Orientation Maneuver ' 00:52:30. 8End Cold-Soak Orientation Maneuver 00:53:26.1

First DPS Burn

Maneuver to Pre-ignition Attitude 03:55:32.0LEM Landing Gear Deployment 03:56:32. 0RCS Ignition 03:57:32.0First DPS Ignition 03:57:38.0First DPS Shutdown 03:58:16. 7

Second DPS Burn

Begin Pre-ignition Orientation Maneuver 04:33:13. 2RCS Ullage Maneuver 04:35:13.2Second DPS Ignition 04:35:19.2Second DPS Shutdown/FITH Staging/First APS

Ignition 04:47:56.2First APS Shutdown 04:48:01.5

42

Table 5-1. Time Sequence ,of-.Events (Continued)

' • • - - . . . Time From LiftoffEvent (hr:min:sec)

Second APS Burn ' '

Begin Pre-ignition Orientation Maneuver 06:09:04. 9RCS Ullage Maneuver - : • 06:11:04. 9Second APS Ignition V.. . - 06:11:07'. 4Second APS Shutdown- ; "" ' . 06:17:42.1

Third APS Burn . • ' • : • ' % : • • - • • . . • • : - . . • :

Begin Pre-ignition Orientation Maneuver ' 07:1.0:46. 1RCS Ullage Maneuver " . 07:12:46. 1Third APS Ignition . '" • '07:12:48.'6Third APS Shutdown "• 07:12:53.9

Spacecraft Orbital Cold-Soak • • • • . ' ' . - . - - • • . .

Third APS Shutdown . . 07:12:53.9Begin Cold-Soak Orientation Maneuver : •• 07:13:03.9End Cold-Soak Orientation Maneuver . 07:13:57.6 ,

Fourth APS Burn • ' • ' " ' • '

Begin Pre-ignition Attitude Maneuver 10:58:53.5.RCS Ullage Maneuver ' ' 11:00:53.5Fourth APS Ignition . - • • - . . . - \ • " i i:QO:56'. 6Fourth APS Shutdown -' 11:01:04.3

Final Orbital Coast

Fourth APS Shutdown • ' • - • • • 11:01:04. 3End of Mission Profile 14:00:00.0

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Table 5-3. Spacecraft Earth Shadow Data

Entrance IntoEarth's Shadow

(Time from Liftoff)(hr :min)

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03:44 •

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6. MSFN COVERAGE DATA

Spacecraft visibility periods for the MSFN stations presented in

Table 3-1 are listed in Table 6-1. Also presented in this table are the

approximate minimum and maximum slant ranges between the spacecraft

and the MSFN station during each visibility period. Table 6-2 presents

the time intervals in which the spacecraft is not visible from a MSFN

station.. The spacecraft is considered visible when the line-of-sight be-

tween the spacecraft and the MSFN station is 85 degrees or less with

respect to the station vertical and the slant range distance is less than

740 nautical miles. The MSFN coverage during major mission events is

illustrated in Figure 6-1.

Volume III presents detailed tracking time history data for the

ground stations available for support of this mission. These data consist

of range, range rate, azimuth angle, azimuth angle rate, elevation angle,

elevation angle rate, and four spacecraft-to-MSFN station look-angles.

Volume III data are presented as a function of time for each of the" ground

stations and annotated for significant mission events.

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Table 6-2. Communication Void Intervals

Void BeginsTime from Liftoff

(hr:min:sec)

OQ:14:08

00:22:38 . . . ' . . .

00:57:45

01:47:08

02:31:39

02:56:04

03:18:19

03:33:14

03:45:26

04:05:06

04:30:34

04:54:13

05:08:25

05:23:59

05:40:58

06:06:27

06:30:17

06:44:38

06:59:27

07:16:57

07:42:51

08:07:31

08:21:17

08:33:11

08:50:27

09:05:16

09:18:43

09:33:51

10:13:20

10:41:13

10:55:25

Void EndsTime from Liftoff

(hr:min:sec)

00:18:09

.00:48:52

01:28:45

02:22:24

02:52:23

03:00:21

03:29:18

03:40:02

03:55:31

04:24:32

04:33:13

05:03:55

05:15:02

05:30:38

06:00:14

06:09:05

06:39:43

06:50:51

07:06:36

07:36:44

07:45:40

08:15:43

08:27:18

08:43:05

08:59:27

09:14:19

09:22:05

10:04:00

10:36:13

10:51:38

10:58:54

VoidDuration(min: sec)

04:01

„. ... ._ -26-16 -

31:00

35:16

20:44

04:17

10:59

06:48

10:05

19:26

03:39

09:42

06:37

06:39

19:16 .

02:38

09:26

06:13

07:09

19:47

02:49

08:12

06:01

09:54

09:00

09:03

03:22

30:09

22:53

10:25

03:29

74

Table 6-2. Communication Void Intervals (Continued)

Void BeginsTime from Liftoff

(hr:min:sec)

11:06:12

11:50:50

12:32:15

13:20:49

' 13:25:18

Void Ends .Time from Liftoff

(hr:min:sec) .

11:40:27 ' • ,

,12;26:35 . " ..

..13:15:56

13:21:12

14:00:00*

Void •Duration

: (min:sec)

.. : 34:15

35:45

43:41

00:23

34:42

* End of mission profile.

75

ASCENT-TO-ORBIT

-LIFTOFF

'f1j!!ill

1 j" i

S-IB INBOARD ENGINES SHUTDOWNS-IB OUTBOARD ENGINES SHUTDOWN

• S-IB + ADAPTER JETTISON/S-IVB IGNITION•» JETTISON BOILERPLATE CSM

(-•-INSERTION INTO ELLI

_ dGRAND TURKfl ^

| BERMUDAGRAND BAHAMA |

CAPE KENNEDY f

BERMUDA SHIP 1

|

III I0:00 0:02 0.04 0:06 0:08 0:10

TIME FROM LIFTOFF (HR:MIN)

0:12 0:14

SLA PETAL DEPLOYMENT

II

TANANARIVE!

0:38 0:39 0:40 0:41


0:42

SPACECRAFT SEPARATION•AHITUDE MANEUVER TOCOLD-SOAK ORIENTATION

KC.1 IU

BEGIN INERTIAL 1ATTITUDE HOLD-.-J

1

MIIIUN

( AUSTRALIA SHIP1

i .'!s SEVER RESTRAINING STRAPSI RCS SHUTDOWN

I I I! 1 i CARNARVON!! 1

1

0:48 0:50 0:52 0:54

TIME FROM LIFTOFF (HR-.MIN)

0:56 0:58

FIRST DPS BURN

-BEGIN PRE-BURN ORIENTATION MANEUVER

iiii1

LANDING GEARDEPLOYMENT ••

j

IRCS IGNITION[K/DPSy ixr i1 1II 1II 1II 1'! I

1 AUSTRALIA SHIP

IGNITION)PS SHUTDOWN

I CARNARVON ||

3:54 3:56 3:58 4:00 4:02


4:04 4K)6

Figure 6-1. MSFN Coverage Summary

76

SECOND DPS BURN/FITH STAGING/FIRST APS BURN

BEGIN PRE-BURN ORIENTATION MANEUVER

/

f^RCS IGNITION f

I l"

P f TEXA

| | WHITE SANDS

J | GUAYMAS |K GOLDSTONE f

I PT.ARGUELLO ]

CALIFORNIA SHIP|

1

!

f BERMUDA SHIP J

f ANTIGUA |BERMUDA |

I GRAND TURK |GRAND BAHAMA J

CAPE KENNEDY J

EGLIN TJ

S J f .APS SHUTDOWN

J \SI FITH STAGING/r APS IGNITION

1

!4:40 4:45

TIME FROM LIFTOFF (HR*1IN)

SECOND APS BURN

XBEG IN PREBURN ORIENTATION MANEUVER

.—RCS IGNITION

Ih——.APS IGNITION

-APS SHUTDOWN

9 r? '

"

1

[ EGLIN :

WHITE SANDS

GUAYMAS1 h GOLDSTONE

1 PT ARGUELLOB CALIFORNIA SHIP 1

I

TEXAS :

11

6:12 6:14 6:16TIME FROM LIFTOFF (HR:MIN)

THIRD APS BURN

L

7:06

j1t

J

1

lj~~

AUSTRALIA SHIP i

7:08 7:10

;— RCS IGNITIONij— APS IGNITIONft-- APS SHUTDOWN

X !T r\^, _ a.T|lTUI>f MANF^VfR JO|J! ! I COLD-SOAK ORIENTATION

*!' CARNARVON \

:!:i i7:12 7:14 7:16 7:18


FOURTH APS BURN

BEGIN PRE-BURN ORIENTATION MANEUVER

^ RCS IGNITION

JJAPS IGNITION

^APS SHUTDOWN

APS

f^A

110:57

1

1 110:58 10:59

"1 1III PT

GOLDSTONE

AGUELLO!' 'CALIFORNIA SHIPin

IIK10

11

1 1 K3i 1 1 ai

1

TIME FROM LIFTOFF (HRJUIN)

Figure 6-1. MSFN Coverage Summary (Continued)

77

7. SUMMARY OF TECHNICAL ACHIEVEMENT

This report contains no innovations or improvements involving new

technology, approaches, methods, or patentable ideas as defined in the

contract's "New Technology and Property Rights in Inventions" clause.

79

REFERENCES

1. "Mission Requirements for Apollo Spacecraft Development MissionAS-206A, " TRW No. 2132-H004-RU-000, 1 December 1965.

2. R. K. Petersburg, "Apollo Mission SA-206A Spacecraft PreliminaryReference Trajectory (U), " TRW No. 3300-H007-RCOOO,1 July 1965. (C)

3. "Apollo AS-206A Trajectory Listing, " furnished MSC by MSFCthrough the Guidance and Performance Subpanel.

4. "Apollo Navigation Working Group, " NASA No. 65-AN-l. 0,5 February 1965.

5. "Operational Support Plan for the Apollo 200 Series Missions, "Prepared by the Flight Control Division/MSC, April 1965.

6. A. Kelemen, "The LEM-1 Mission Capability Report NASA MissionAS-206A (U), " No. LED-540-41, 1 November 1965. (C)

7. Minutes of Flight Operations Plan, Meetings 1 through 12.

8. "Apollo Mission Data Specification C Apollo-Saturn 206A (U), "TRW No. 2131-6002-TCOOO, 5 November 1965. (C)

9. "Launch Vehicle Reference Trajectory, " MSFC, to be published.

10. "Lifetime of Near Earth Satellites in Circular or Elliptical Orbits, "Memorandum for Record, OFO (JCB:jec), 1 3 September 1 963. (C)

11. "Apollo Guidance and Navigation, G and N System Operations PlanMission AS-206, " MIT Instrumentation Laboratory No. R-527,November 1965. (C)

81

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