Post Launch Report for Apollo Mission a-102

8/6/2019 Post Launch Report for Apollo Mission a-102

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UNCLASSIFIED

POSTLAUNCH REPORT FOR

APOLLO MISSION A-102 ( e >(BP- 151

Approved fo r Di s t r i bu t i on :

A D r . #oseph @’.&eap o l l ~ # pacecraft Program Office

NATIONAL AERONAUTICS AND SPACE ADM IN I S I TRAT ION

MA- SPACECRAFT CENTER

HOUSTON, TEXAS

October 10, 1964

UNCLASS IF IED



U N C L A S S l F l E D

CONTENTS

Page

1.0

2 .0

3 * 0

4.0

E

J .0

6.0

ABBREVIATIONS AND SYMBOLS . . . . . . . . . . . . . .. . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .bbrevietionsSymbols

'IIABLES . . . . . . . . . . . . . . . . . . . . . . .

SUMMARY . . . . . . . . . . . . . . . . . . . . . . .INTRODUCTION . . . . . . . . . . . . . . . . . . . .FLIGHT TRAJECTORIES . . . . . . . . . . . . . . . . .

SPACECRAFT DESCRIPTION AND PERFORMANCE . . . . . . .

4.14.2

4.3

4.44.54.6

4 .74.8

4.94.104.114.12

Spacecraft Description . . . . . . . . . . . .Instrumentation . . . . . . . . . . . . . . . .Electrical Power and Sequential . . . . . . . .Communications . . . . . . . . . . . . . . . .Pyrotechnic Devices . . . . . . . . . . . . . .Launch-Escape Subsystem Propulsion . . . . . .Structures . . . . . . . . . . . . . . . . . .React ion C o n t r o l Subsystem . . . . . . . . . .Acoustics . . . . . . . . . . . . . . . . . . .Heat Protection . . . . . . . . . . . . . . . .Aerothermodynamics . . . . . . . . . . . . . .Equipment Cooling . . . . . . . . . . . . . .

SA-7 LAUNCH-VEHICLE DESCRIPTION AND PERFORMANCE . . .

5.1 Description . . . . . . . . . . . . . . . . . .5.2 Preliminary Flight Performance . . . . . . . .CONCLUDING RE3lARKS . . . . . . . . . . . . . . . . .

iii

iiiv

vi1

ix

xx i

1-1

2-1

3-1

4-'1

4-14-144-28

4-334-394-454-524-12:>

4- 72

4- 7f')4.18. .4-21:

7-1

5-1

5-1

6-1

U N CLASS1F I ED



ii

7.0

8.0

9.0

10.0

U N C L A S S 1 F lED

APPEXDIX A

7.1 Prelaunch Operations . . . . . . . . . . . . . . 7-17.2 Launch Operations . . . . . . . . . . . . . . . 7-10

7.3 Range Operations . . . . . . . . . . . . . . . . 7-157.4 Data Coverage and Availability . . . . . . . . . 7-197.5 Telemetry Tape Selection and Verification . . . 7-38

A P P E N D I X B . . . . . . . . . . . . . . . . . . . . . . 8-1

APPENDIX C

9.1 C-Band Beacon Anomaly . . . . . . . . . . . . . 9-19.2 Loss of Telemetry Measurement, SR5877T . . . . 9-59 . 3

(SA0553R) . . . . . . . . . . . . . . . . . . 9-6Loss o f H e a t Flux Data F r o m C a l o r i m e t e r 13

R E F E R E X C E S e . . - . . . . . . . . . 10-1

U N C LA SSI F I ED



UNCLASS I F lED

ABBREVIATIONS AND SYMBOLS

ii i

Abbr e v i a ti ns

A I

BP

CCW

CM

cw

D.A.D.

ECS

ETR

FM

FO

g . e . t .

G . m . t .

G SE

IE CO

IF

D I G

ru

KSC

LE S

LOV

l o x

MCC

adap te r and in se r t

b o i l e r p l a t e

counterclockwise

command module

clockwi s

double ampli tude de f le c t i on

environmental control system

Eastern T e s t Range

frequency modulation

Flor ida Opera t ions

ground elapsed time

Greenwich mean t i m e

ground support equipment

inboard engine cutoff ( %I s tage)

intermediate frequency

in ter range ins t rumenta t ion group

inst rument un i t

John F. Kennedy Space Center

launch escape subsystem

loss of v e h i c l e

l iquid oxygen

Mission Control Center

UNCLASS IF1ED



i v

MILA

MSC

MSFC

OECO

OTP

-P

+P

PAM

PIA

m

PLIM

R and D

RC S

SA

SM

SPL

sr

s-IV

T

TCV

mvco

VRF

VSWR

U N C L A S S I F I E D

Merritt Island Launch Area

Manned Spacecraft Center

Marshal l Space Fl igh t Cen t e r

outboard engine cutoff ($1 stage)

Operat ional Test Procedure

nega t i ve p i t ch

p o s i t i v e p i t c h

pulse amplitude modulation

pre - ins t a l l a t i on accep t ance

pul se repe t i t i on f requency

post launch ins tnunenta t ion message

Research and Development

reac t ion cont rol subsys tem

Saturn- Apollo

s e r v i c e module

sound p r e s s u r e l e v e l

Saturn launch vehicle , f i r s t s tage

Saturn launch vehi cle , second stag e

launch t ime

thermal cont rol va lve

te lemet ry

v o l ta g e c o n t r o l l e d o s c i l l a t o r

very high frequency

vol tag e standing wzve r a t i o




V


Symbols

e At

expansion area ra t i o , c ros s - sec t i ona l area a t some point

i n nozzl e d iv ided by c ross - sec ti ona l a rea a t t h roa t

m a x i m u m body diameter

thrus t , l b f

g r a v i t a t i o n a l c o n s t an t

moment of i ne r t i a a round t he X-axis, s l u g - f t2

D

F

2moment of i n e r t i a around t he Y - a x i s , s l u g - f t

IYY

2moment of i n e r t i a around Z - a x i s , s lug- f t

c h a r a c t e r i s t i c l e n g t h

Mach number

pre ssure, l b / s q i n .

dynamic pressure, lb/sq f t

maximum dynamic pressure, l b / s q f t%ax

2h e a t f lux, Btu/ft /sec

Reynolds number, based on m a x i m u m body diameter DReD

RMS root mean square

T launch t i m e , sec

X l ong i tud ina l ax i s o f t h e spacecraf t and launch vehic le

xAlongi t udina l loca t io n, re fe renced t o t he s2acecra f t , i n .

( f i g . 4.1-3)

xClo ng itu di na l loca tio n, refer enced t o th e command module, in .

( f i g . 4.1-3)

xLlongi t udina l loca t io n, re fe renced t o the launch-escape

subsystem, i n . ( f i g . 4.1-3)




vi U NCLASS IF1ED

xLv

xS

Y

Z

l ong i tud in a l l oca t ion , r e ferenced t o the launch-vehicle

S I t ag e , i n . ( f i g . 4.1-3)

l ong i tud ina l l oca t i on , r e ferenced t o t he se rv i ce module, i n .

( f i g . 4.1-3)

plane of th e Y-axis passes through the X - a x i s and i s per-pendicula r t o t he p l ane of t he Z -axi s, i n . ( f i g . 4.1-2)

plane of t he Z - a x i s passes through the X-axis and through

t h e c e n t e r of t h e CM hatch and of f i n s I and I11 of t h e

SA-7 l aunch veh ic l e , i n . ( f i g . 4.1-2)

a angle of a t tack, deg

aq product of angle of attack and dynamic pressure,

(deg) ( lb /sq ft)

a &ball differential pressure

F l i g h t t e s t symbols, Saturn-Apollo: Saturn symbol within

Apollo symbol. Basic symbols: Greek and Roman.




U NCLASS IF IED

TABLES

vi

Tzble P a g e

2-3

3-4

APOUO SPACECRAFT FLIGHT HISTORY . . . . . . . . . . ..0-1

3.0-1 M i SSI O N EVENT TIMES . . . . . . . . . . . . . . . . .

COMPARISON OF SA-7 PLANNED AND ACTUAL TRAJECTORY

PARAMETERS . . . . . . . . . . . . . . . . . . . . ..0-11

3-5

4- 4 '

4-5

4 . 1 - 1

4.1- 1

4 . 2 - 1

WEIGHT COMPARISON OF BP-13 AND BP-15 SPACECRAFT. . . .

SPACECFAFT BP-15 MASSCHARACTERISTICS. . . . . . . . .

APOLLO MI SS IO N A -1 0 2 MEASUREMENT REQUIREMENTS

S W Y . . . . . . . . . . . . . . . . . . . . . . . 4-17'

4.2-11 FLIGHT EQUIPMENT FOR BP-15 SPACECRAFTINSTRUMENTATION SUBSYSTEM. . . . . . . . . . . . . . 4-15,

4-41.5-1

4.6-1

ASSIGNMENT OF PYROTECHNIC DEVICES. . . . . . . . . . .

BP-15 SPACECWXI LAUNCH-ESCAPE PROPULSIONSUBSYSTEM MOTOR PREDICTED PERFORMA-NCE. . . . . . . . 4-4G

4.8-1 COMPARISON OF BP-15 SPACECRAET INSTRUMENTEDRCS CHAMBERS WITH PROTOTYPE RCS CHAMBER.

. . . . . .4.8 -11 COMPARISON OF CALCULATED AND RECORDED MAXIMUM

TEMPERATURES. SERVICE M0DUL;E RCS QUAD A.BP-15 SPACECRAFT . . . . . . . . . . . . . . . . . .

MONITORED EQUIPMENT COOLING SUBSYSTEM

P A R A M E T E R S . . . . . . . . . . . . . . . . . . . . .

4-135

4.12-14 - 2 1 ~ 4

4.12-11 EQUIPNl3NT COOLING SUBSYSTEM PARAWERSAT UMBILICAL DISCONNECT ( T - 1 8 SE C ) . . . . . . . . . 4-21.3

7-1-1

7-itj

7-21-

7-3-1

7.3-11

7.4-1

TELEMETRY COVERAGE . . . . . . . . . . . . . . . . . .

C-BAND RADAR COVERAGE. . . . . . . . . . . . . . . . .

DATA AVAILABILTTY. . . . . . . . . . . . . . . . . . .




v i i i U NCLASS IF IED

Table Page

7.4-11 EN GI NE ER IN G SEQUEXI"T CAMERA DATA . . . . . . . . . . 7-30

7.4-111 COMPARISON OF INSTRUMEITIATION MEASUREMENTSUSED ON BP-13 AND BP-15 SPACECRAFT . . . . . . . . . 7-33

7 . 4 - I V LAUNCH-VEHICLE DATA PROCESSED FOR hEC-HOUSTONA T K F C . . . . . . . . . . . . . . . . . . . . . . . 7-34

. . . . . . . . . . . . . . . .. 4 - V WFC D A T A A V A I L A B I L W 7-35

8.1-1 MEASTJREMENTIST FOR BOILERPLATE 1 5 SPACECRAFT . . . . 8-2

8.2-1OPERATIONAL

TESTPROCEDURES FOR APOL LO

BP-15SPAC ECRA FT AT CONTRACTOR'S MANUFACTURINGF A C I L I T Y . . . . . . . . . . . . . . . . . . . . . . 8-11

8.2-11 OPERATIONAL TEST PROCEDURE FOR APOLLO BP-15SPACECFKFI' AT FLORIDA OPERATI ONS . . . . . . . . . . 8-12

9.1-1 SUMMARY OF SEIXCTED RADAR m PO Fi T S ON

ORBITAL PASSES 1AND 2 . . . . . . . . . . . . . . . 9-3




Y’g1 re

2.0-1

2.0-2

3.0-1

3.0-2

3.0-3

3.0-4


FIGURES

ix

Page

Saturn-Apollo space vehi cle f o r mission A - 1 0 2a t l i f t - o f f . . . . . . . . . . . . . . . . . . . . . 2-4

Sequence of major events f o r Apollo mission A-102 . . . 2-5

Ground track for th e Apollo A-102 o r b i t a l mission. . . . . . . . .o r t h e f i r s t t h r ee o r b i t a l p as se s 3-7 .

Al t i tude - long i tude p ro f i l e f o r Apollo missionA-102 f o r t h e f i r s t t h r e e o r b i t a l p as s es . . . . . . 3-8

Tim e h i s t o r i e s of t r a j ec to ry paramete rs for t heApo llo mis sio n A-102 laun ch pha se

( a ) Alt i tude and range . . . . . . . . . . . . . . . . 3-9( b ) Space-f ixed vel oci ty and. f l ig ht -pa th angl e . . . . 3-10(c ) Ear th- f ixed ve loc i ty and f l i gh t -pa t h angle . . . . 3-11( d ) Dynamic pressure and Mach number . . . . . . . . . 3-12

( e ) Longi tud inal acce lera t ion a long spacecraf tX-axis . . . . . . . . . . . . . . . . . . . . . 3-13

Time histor ies of t ra jec tory parameters for Apollo

mission A - 1 0 2 for f i r s t t h re e pa ss es of o r b i t a l

phase

( a ) Lat i tude , longi tude , and a l t i tu de . . . . . . . . 3-14(b) Space-fixed ve lo ci ty and flight-path angle . . . . 3-13

4 . 1 - 1

4.1-2

4.1-3

Apollo BP-15 spacecraf t . . . . . . . . . . . . . . . . 4-6

Y- and Z-axes and ang ula r coor din ate system =sed

for designating locations within the BP.-15

. . . . . . . . . . . . . . . . . . . . .pacecraf t 4-7

X-axis systems used fo r desi gnat ing long itu din al

loca t ions o f BP-13 spacec ra f t . . . . . . . . . . . . 4-E

Launch escape subsystem for BP-15 spacec ra f t . . . . . 4-5

Command module in t e r i o r equipment la yo ut f o rBP-15 spac ecraf t (view through hatch ) . . . . . . . . 4-10




X

Figure

4.1-6

4.1-7

4.1-8

4.2-1

4.2-2

4.2-3

4.2-4

4.2-5

4.2-6

4.2-7

4.2-8

4.3-1

4.3-2

4.3-3

4.4-1

4.4-2


Page

Command module interior equipment layout(view to right of hatch) f o r BP-15 spacecraft. . . 4-11

Command nodule exterior of BP-15 spacecraft . . . . 4-12

Cutaway view of Bp-15 spacecraft service module,insert, and adapter . . , . . . . . . . . . . . . 4-13

Instrumentation and comunications subsystems onBP-13 spacecraft . . . . . . . . . . . . . . . . . 4-20

Locations of linear acceleration transducers forBP-15 spacecraft . . . . , . . , . . . . . . . . , 4-21

Stra in-gage locations on BP-15 spacec ra f t . . . . . 4-22

Fluctuating-pressure transducer locations onBP-15 spacecraft . . . . . . . . . . . . . . . . . 4-23

Locations of conical surface pressure transducerson the BP-12 spacecraft , , , . . . . . . . . . . 4-24

Locations of heat-flux calorimeter bodytemperature measurements on BP-15 spacecraft . . . 4-25

Locations of transducers on BP-15 spacecraft . . . . 4-26

Command module interior showing cable shieldingto prevent EMI in the BP-15 spacecraft . . . . . . 4-27

Electrical power subsystem for BP-13spacecraft . . . . . . . . . . . . . . . . . . . . 4-30

Electrical power subsystem components forBP-15 spacecraft . . . . . . . . . . . . . . . . . 4-31

Launch escape sequencer subsystem for BP-15spacecraft . . . . . . . . , . . . . . . . . . . . 4-32

bcation of telemetry transmitters and C-band

transponders on i3P-15 spacecraft . , , . . . . . . 4-35

bcation of telemetry omniantenna on commandmodule of Bp-15 spacecraft . . . . . . . . . . . . 4-36




U N C L A S S I F I E D x i

PageFigure

4.4-3

4.4-4

4.5-1

4.5-2

4.5-3

4.6-1

4.6-2

4.6-3

4.6-4

4.6-5

4.6-6

4.7-1

4.7-2

4.7-3

4.7-4

4.7-5

BP-15 spacecraft C-band beacon antenna locations

on the se rv i ce module . . . . . . . . . . . . . . . 4-37

C-band transponder block diagram for BF-15s p a c e c r a f t . . . . . . . . . . . . . . . . . . . , 4-38

Location of i g n i t e r c a r t ri d g es i n BP-15spacec ra f t . . . . . . . . . . . . . . . . . . . . 4-42

BP-15 launch escape tower explosive bolti n s t a l l a t i o n . . . . . . . . . . . . . . . . . . . 4-43

BP-15 s p a c e c r a f t LES explosive bo l t d e t a i l . . . . . 4-44

BP-15 spacecraf t launch escape subsystem . . . . . . 4-49

EP-15 I;ES launch escape motor . . . . . . . . . . . 4-50

BP-15 LES launch-escape motor pre dic ted th ru st

( i n vacuum a t 70" F) . . . . . . . . . - . . . . . 4-51

BP-15 spacec ra f t IXS p i t ch con t ro l mo to r . . . . . . 4-52

BP-15 IXS p i t c h co n t ro l motor p r ed ic ted t h r us t

( i n vacuum a t 70" F) . . . . . . . . . . . . . . . 4-53

BP-13 spacec ra f t U S tower j e t t i s on mo to r . . . . . . 4-54

Apollo BP-13 spacec ra f t launch escape subsystem

s t r u c t u r e . . . . . . . . . . . . , . . . . . . . . 4-67

Detail of command module-service module interface

f o r BP-15 s p a c e c r a f t . . . . . . . . . . . . . . . 4-68

Rawinsonde atmospheric wind data at Cape Kennedy,

Fla . , Sep t . 17-18, 1964 . . . . . . . . . . . . . . 4-69

Comparison of p red ic te d t o f l i g h t measuredangles of a t t ac k and c q , . . . . . . . . . . . . . 4-70

S ta t i c p ressure coe f f i c ie n t over the command

module conical surface on Bp-15 s p a c e c r a f t

(a Circumferential location, approximate* 90". . . 4-71(b) Circumferen t ia l loca t ion , 180' . . . . . . . . . 4-72

( c ) C i r cumfe ren t i a l l oca t ion , 3570 . . . . . . . . . 4-73

UNCLASSIF IED



U NCLASS F I EDi i

Pageigure

4.7-6 Command module s t a t i c pr ess ure co ef fi ci en ts f o rBP-15 spac ecraf t f l i gh t da ta compared t o windtunnel da ta ( reference 6) . . . . . . . . . . . . . 4-74

Pressure venting scheme for BP-15 spacecraftser vic e module, in se rt , and adap tercompartment . . . . . . . . . . . . . . . . . . . .

4.7-7

4-75

4-764.7-8

4.7-9

BP-15 sp ac ec ra ft ser vi ce module in t er na l

pressure . . . . . . . . . . . . . . . . . . . . .

ins trumen t un i t in te r na l p ressure ( f l i gh t

BP-15 sp ac ec ra ft se rv ic e module and laun ch-v ehic le

measured d a t a ) . . . . . . . . . . . . . . . . . . 4-77

Comparison of BP-15 t o BP-13 command moduleinstrumentation compartment internalpressures . . . . . . . . . . . . . . . . . . . . .

4.7-10

4-78

4-79Poss ib le a i r leakage path caused by f r a c tu r e

of explosive bo lt . . . . . . . . . . . . . . . . .Fl igh t measured acceleration (BP-15 spacecraft )

4-7-11

4.7-12

(a ) Lunch-e scape subsystem a t &-bal li n t e r f a c e . . . . . . . . . . . . . . . . . .

(b ) Command module . . . . . . . . . . . . . . . . .4-80

4-82

To ta l ax ia l fo rce a t in te r face of BP-15 spacecraft

ad ap te r and Saturn SA-7 inst rum ent un it(XA722) . . . . . . . . . . . . . . . . . . . . . .

Power spectral distribution command module X-axisaccelera t ion , E-15 spacecraf t

4-84

( a ) T+48 seconds . . . . . . . . . . . . . . . . . .(b ) T+73 seconds . . . . . . . . . . . . . . . . . .

4-834-86

Power s pec tra l d is t r i but ion , Y- and Z-axesaccelerations, BP-15 spacecraft

4.7-15

4-874-88

(a) Tower, Y-axis, T+48 seconds,(b ) Tower, Z-axis, T+48 seconds,

I A O O l l A . . . . . .IAOO12A . . . . . .

U NCLASS F ED



U N C L A S S I F I E D x i i i

Figure Page

(e)

(d)

Command module. Y.axis. T+48 seconds.

Command module. Z.axis. ~+48econds.CAOOO5A . . . . . . . . . . . . . . . . . . . 4-89

CAOOO7A . . . . . . . . . . . . . . . . . . . 4-90

4.7-16 Development view of BP-15 spacecraft service

module. insert. and adapter wall showinginstrument locations . . . . . . . . . . . . . . . 4-91

4.7-17 RMS history of service module radial vibration

(a) Sensor SA0086D . . . . . . . . . . . . . . . . 4-92(b) Sensor SA0087D 4-93(c) Sensor SA0088D . . . . . . . . . . . . . . . . 4-94

. . . . . . . . . . . . . . . .

4-7-18 RMS of fluctuating pressure over BP-15 spacecraft

Sensor SAO162PSensor SAO163PSensor SAO164P

Sensor SAO165P

Sensor SA0166PSensor ~ ~ 0 1 6 7 ~Sensor SAO168PSensor SAO169PSensor SAO17OP

Sensor SA0171PSensor SAO172PSensor AA0173P

Sensor AA0174P

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .. . . . . . . . . . . . . . . .. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .4.7-19 Power spectral distribution service module radial

vibration. BP-15 spacecraft

4-95

4-97

4-99

4-96

4-98

4-1004-1014-102

4-103

4- 044-1054-106

4-107

(a) SA0086D. T+49 seconds . . . . . . . . . . . . . 4-108(b) SA0088D. T+49 seconds . . . . . . . . . . . . . 4-109

4.7-20 Power spectral distribution of BP-13 servicemodule radial vibration at T+5O seconds . . . . . 4-110

4.7-21 Parer spectral distribution of BP-15 servicemodule. strain (SA2121S) . . . . . . . . . . . . . 4-111




xiv U N C L A S S IF1ED

Figure

4.7-22 Time h is t or y of BP-15 sp ac ec raf t ser vi ce modules t r a i n

Page

(a) Sensor SA212OS . . . , . . . . . . . . . . . . .( b ) Sensor SA2121S . . . . , . . . . . . . . . . . .

4-1124-114

4.7-23 RMS of BP-15 se rv ic e module s t r a i n

(a ) ~A2120S . . , . . . . . . . . . . . . . . . . .(b) SA2121S . . . . , . . . . . . . . . . . . . . .

4-1164-117

4.7-24 Time h i s t o r y o f BP-15 se rvi ce module qua si s teadys ta te s t r a i n - s~2120s averaged)

(a) SA2120S . . . . . . .. . . . . . . . . . . . . . 4-118(b) SA2121S . . . . . . . . . . . . . . . . . . 4-120

4.7-25’ Time hi st or y of BP-15 ada pter s t r a i n

(a) AAO195S and ~ ~ 0 1 9 6 s . . . . . . . . , , . . . 4-122( b ) AA0197S and ~ 0 1 9 8 s . . . . . . . . . . . . . . 4-123

4.7-26 BP-15 spacec raf t adapter load from s t ra in gage

data a t ~ ~ 7 3 6. . . . . . . . . . . . . . . . . 4-124

4.7-27 Comparison of f l i g h t nois e data t o d es i gn

environment

( a )

( b )

( c )

(d )

External-forward of XA910 t o XA1017.75 . . . . .S M RCS engine area XAg10 t o XA959 . . . . . . .Ekternal XA838 t o XA910 . , , . . . . . . . . .Ekternal XA722 t o XA838 . . . . . . , . . . . .

4-125

4-126

4-127

4-128

4.8-1 Loca tion f o r s er vi ce module RCS quads onRP-15 spacec ra f t . . , . . . . . . . . . . . . . . 4-136

4.8-2 Service module RCS quad A on B-15 spacec ra f t , . . . 4-137

4.8-3 I n t e r i o r view of s e rv ic e module RCS quad A for

BP-15 spacecraft . , . . . . , . . . . . . . . . . 4-138

4.8-4 Spacecraf t BP-13 quad A and prototype RCSchamber design . . , . . . , . , . . . . , . , . . 4-139

UNCL AS S IF IE D



U NCLASS IF I ED xv

F-lgure Page

j1.8-5 Method of mounting service module RC S engine to quad

housing and engine supporting bracket on BP-15

spacecraft . . . . . . . . . . . . . . . . . . . . . . 4-140

quads B, C , D of BP-15 spacecraft . . . . . . . . . . 4-141

spacecraft . . . . . . . . . . . . . . . . . . . . . . 4-14:?

11.8-6 Design of dummy service module RC S chambers used in

MSC coordinate axes and notation system for Apollo

boilerplate and airframe, manned and unmanned4.8-7

4.8-8 Service module RC S package instrumentation locationsfor clockwise and counterclockwise r o l l engines,engine supporting bracket, and quad housing roofon BP-15 spacecraft . . . . . . . . . . . . . . . . . 4-145

4.8-9 Service module R C S package instrumentation locationsfor +P and -P engines, engine supporting bracket,

and quad housing roof on BP-15 spacecraft . . . . . . 4-141-

4.8-10 Mounting of thermocouples on the service module RCS

engines on BP-15 spacecraft . . . . . . . . . . . . .4.8-1.1 Service module RCS accelerometers mounted in the

clockwise engine nozzle on BP-15 spacecraft . . . . .4.8-12 Temperatures measured on the positive pitch engine

and housing of the service module RCS quad A of

the BP-15 spacecraft

(a) -25 to 325 seconds . . . . . . . . . . . . . . . .

( c ) 675 to 850 seconds . . . . . . . . . . . . . . . .Temperatures measured on the counterclockwise roll

( b ) 327 to 673 seconds . . . . . . . . . . . . . . . .

4.8-13engine and housing of the service module R C S

quad A o f the BP-15 spacecraft

(a) -25 to 325 seconds . . . . . . . . . . . . . . . .(b) 323 to 675 seconds . . . . . . . . . . . . . . . .( e ) 675 to 850 seconds . . . . . . . . . . . . . . . .

Temperatures measured on the negative pitch engineand housing of the service module RC S quad A ofthe BP-15 spacecraft

4-8-14

(a) -25 to 325 seconds . . . . . . . . . . . . . . . .

4-14:,

4-141;

4-1504- 314-15?




xvi

Figure

4.8-15

4.8-16

4.8-17

4.8-18

4.8-19

4-8-20


(b) 325 to 675 seconds . . . . . . . . . . . . . . .(c) 675 to 850 seconds . . . . . . . . . . . . . . .Temperature measured on the engine supporting

bracket of the service module RCS quad Aof the BP-15 spacecraft

(a) -25 to 325 seconds . . . . . . . . . . . . . . .(b) 325 to 675 seconds . . . . . . . . . . . . . . .(c) 675 to 850 seconds . . . . . . . . . . . . . . .Temperature measured on the underside of thequad housing roof of the service module

RCS quad A of the EP-15 spacecraft

(a) -25 to 325 seconds . . . . . . . . . . . . . . .(b) 325 to 675 seconds . . . . . . . . . . . . . . .(c) 675 to 825 seconds . . . . . . . . . . . . . . .

Temperatures measured on the +P, counterclockwise,

and -P engine injector heads and on the enginesupporting bracket of the service module RCS

quad A of the E€'-15 spacecraft

(a) 0 to 350 seconds . . . . . . . . . . . . . . . .( b ) 350 to 700 seconds . . . . . . . . . . . . . . .(c) 700 to 850 seconds . . . . . . . . . . . . . . .Temperatures measured on the +P, counterclockwise,

and -P engine housings and on the quad housingroof of the service module RCS quad A of theBp-15 spacecraft

(a) 0 o 350 seconds . . . . . . . . . . . . . . . .(b) 350 to TOO seconds . . . . . . . . . . . . . . .(c) 700 to 850 seconds . . . . . . . . . . . . . . .X-axis vibration measured in the clockwise roll

engine nozzle of the service module RCS

quad A of the,BP-15 pacecraft . . . . . . . . . .Perpendicular a x i s vibration measured in the

clockwise roll engine nozzle of the servicemodule RCS quad A of the BP-15 spacecraft . . . . .

Page

4-1344-155

4-1564-1574-158

4-1594-1604-161

4- 624-163

4- 64

4-165

4-1674- 66

4-168

4-169

U NCLASS IF EO



4.8-22

4.9-1

4 . “-2

L L . 3-3

. ( . J b

4.10-1

4.10-2

4. o-3

4. l-1

4.11-2

U NCLASS F1ED

Di gi ta l spectrum est imat ion of X-axis vi br at io n

over the t i m e pe ri od T+48 t o Ti-50 secondsmeasured in the clockwise roll engine

nozz le of t h e se rv ic e module RCS quad A

of t h e RE’-15 spacec ra f t . . . . . . . . . . . . . .

Digi tal spestrum est imat ion of t h e pe rpendicu la r

ax is v ib ra t io n over t h e period Ti-48.01 t o

Ti-30.01 seconds measured in the clockwise

r o l l engine nozzle of t h e se rv ic e module

RCS quad.A of t h e BP-15 spacec ra f t . . . . . . . .

Overall sound pressure l e v e l time h i s t o r y fromT-10 t o T + 1 1 0 seconds for B P - 15 spacec ra f t . . . .

One-third octave band an aly si s of BP-15 space -

craft launch-vehicle sound pre ssure l eve l a t

T + 1 Eecsnd . . . . . . . . . . . . . . . . . . . .Comparison of BP-13 and BP-15 leunch-vehicle sound

pressure levels from T-2 t o ~8 seconds . . . . . .

One-third octave band analysis of BP-1, aerodynaaic

sound pressure le ve ls a t W53 seconds . . . . . . .

Command module he a t pr o te c ti on f o r BP-15spacec ra f t

. . . . . . . . . . . . . . . . . . . .Launch-escape t o w e r t empera tu re t ransducer

locations on BP-15 spacec ra f t . . . . . . . . . . .Bond-line LES tower temperatures measured during

f l i g h t of BP-15 spacec ra f t

( a ) Tower temperatures 1, 2 and 8 . . . . . . . . .(b ) Tower temp eratu res 4 , 5 and 7 . . . . . . . . .( c ) TDwer temperatures 3 and 6 . . . . . . . . . . .Top view of BP-15 spacecraft command module

showing calor imet er loca t ion s . . . . . . . . . . .Development view of m-15 pacec ra f t s e rv i ce

module, i n se r t , and adapter compartment

showing ca lor imeter loca t ions . . . . . . . . . . .

x v i i

Page

4-170

4-171

4-174

4- 180

4-181

4-182

4- 834-184

4-189

4-19




xviii

Figure

4.11-3

4.11-4

4.11-5

4. 11.6

4.11-7

4.11-8

4.12-1

U NCLASS IFI ED

hunch configuration environment in terms of

Mach number (M ) and Reynolds number (ReD)

f o r BP-15 spacecraft . . . . . . . . . . . . . . .Angle of at ta ck hist ory fo r BP-15 space craft

(Q-ba l l data) . . . . . . . . . . . . . . . . . .Heating rates measured on BP-15 spacecraft

command module

Calorimeters 1. 5. and 10 .Calorimeters 6. 7. and 8 .Calorimeters 2. 4. and 11 .Calorimeters 3 . 9. and I 2 .Calorimeters 1. 2. and 3 .Calorimeters 4. 5 . and 9 .Calorimeters 10. 11. and 12Calorimeters 1. 2. and 3 .Calorimeters 11 and 12 . .Calorimeters 4. 5 . 6 and 9Calorimeters 3. 9 and 12 .Calorimeters 2. 4 and 11 .Calorimeters 1. 5 and 10 .Calorimeters 6. 7 and 8 . .

. . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . .. . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . .

Comparisons of BP-13 and BP-15 s pa cec ra f t ra te

( a ) BP-13 and BP-15 cal or im et er 3( b ) BP-13 and BP-15 calorimeter 5( c ) BP-13 and Bp-15 calor imeter 10 . . . . . . . .(a) BP-13 and BP-15 calorimeter 16 . . . . . . . .

. . . . . . . . .

. . . . . . . . .

Heating rates measured on BP-15 spacecraf t

service module

(a) Calorimeters 14 and 16 . . . . . . . . . . . .(b) Calorimeters 15 and 19 . . . . . . . . . . . .(c) CalorLmeters 17 and 18 . . . . . . . . . . . .(d) Calorimeters 13 and 20

. . . . . . . . . . . .Comparison of cal ori met er body temperatures a tloca t ions 13 and 20 . . . . . . . . . . . . . . .

Environmental control subsystem schematic forBP-15 spacecraft . . . . . . . . . . . . . . . . .

4- 91

4- 192

4-1934-194

4-195

4-1964-1974-984-199

4- 004- 2014- 202

4- 2024- 203

4- 203

4-200

4-2044-2044-20?4-20?

4-206

4-2084-207 .

4-209

4-210

4-216

U NCLASS I F IED



U N C L A S S I F I E D x i x

PageFigure

4.12-2

4.12-3

4.12-4

5.1-1

7.1-1

7.1-2

7.1-3

7.1-4

7.1-5

7.2-1

7.4-1

7.4-2

9.1-1

9 .3 -1

Equipment cooling prelaunch data f o r Bp-1.5spacec ra f t . . . . . . . . . . . . . . . . . . . . . 4-217

S ec ti on al view of coolant-pump assembly forBP-15 spacecraft . . . . . . . . . . . . . . . . . . 4-218

Command module cabin pressure for BP-15spacec ra f t . . . . . . . . . . . . . . . . . . . . . 4-219

Apollo missi on A-102 space v eh ic le showingcutaway views of launch vehicle . . . . . . . . . . 5-4

Spa cecr aft BP-15 schedul e milesto nes a t AT0

Downey . . . . . . . . . . . . . . . . . . . . . . . 7-5

Sp ace cra ft BP-15 s che dule mile sto nes a t MSCFlorida Operations . . . . . . . . . . . . . . . . . 7-6

BP-15 spacecraft CM SM being stacked on adapters e c t i o n i n Hangar AF, Cape Kennedy, Flo r ida . . . . 7-7

BP-l'j spacecraf t be ing l i f t e d t o mate wi th the

SA-7 lau nch ve hi cl e a t Launch Complex 37B,. . . . . . . . . . . . . . .ape Kennedy, Florida 7-8

U S b ei ng l i f t e d f o r ma ti ng t o t h e s p a c e c r a ft

a t Launch Complex 37B, Cape Kennedy, Flo r ida . . . . 7-9

Apollo mission A-102 countdown ac t i v i t i e s

. . . . . . . . . . .a) 'T-1 ay, September 17, 1964 7-12. . . . . . . . .b) Launch day, September 18, 1964 7-13. . . . . . . . .c ) Launch day, September 18, 1964 7-14

Engineering seq uen tia l t ra cki ng camera loca tio ns

. . . . . . . . . . . . . .o r Apollo mission A-102 7-36

Ehgineering seq uen tia l f ixe d camera loc ati ons

. . . . . . . . . . . . . .o r Apollo mission A-102 7-37

BP-15 s pace craf t t ransponder in te rr og at io n and

. . . . . . . . . . . . . . . . . . . . 9-4e s p o n s e . ,

BP-15 spacec ra f t ca lo r ime te r . . . . . . . . . . . . . 9-8




Preceding page blank

UNCLASSIF IED xxi

CONTRIBUTORS

Spacecraft Description: T. Dalton, W. H. Waln, D. G. Nichols,A. L. Branscomb, A. L. Cave, V. V. Sheeley, W. F. Edson.

Instrumentation, Electrical, Sequential, and Communications:J. F. Rutherford, J. B. Pennington, J. M. Eller, P. Katalansky,F. M. Groark, R. Stevens, G. Manning, R . E. McCoy, R. F. Jones,T. G. Broughton, C. P. Lassetter, A. E. Warner.

Pyrotechnics and Propulsion: W. H. Simmons, P. Davis, H. E. Heilman,L. J. Haywood, E. A. Timmons, J. 0. Payne.

Structures and Acoustics: W. M. West, B. V. Zuber, S. P. Weiss,W. F. Rogers, B. 0. French.

Reaction Control Subsystem: N. H. Chaffee

Heat Protection and Aerothermodynamics: J. E. Pavlosky,N. W. Willis, R. M. Raper.

Equipment Cooling: D. F. Hughes, J. R. Hiers, S. P. Moody.

Trajectory and Launch Vehicle: D. J. Incerto, W. P. Beal,J. L. Wells, M. Cassetti, R. D. Nelson, R. Teasley, G. Emanuel.

Mission Operations, Data, and Tape Selection: F. Peters,

J. M. Gerding, R. E. Reyes, K. Sorey, D. T. Jensen, D. M. Goldenbaum,T. J. Grace, S. A. DeMars, F. B. Blanton, M. V. Britt, E. C. Wang, .R. C. Shirley, C. E. R o t h , J. M. E l l e r , R. H. T alb e r t , W. T. Lauten.

Anomalies: C. J. Appelberg, E. Rangel, G. Zilling.

Editing, Summary, and Concluding Remarks: J. D. Lobb,C. T. Stewart, F. Peters, J. M. Gerding, E. 0. Zeitler, R. E. McKann,W. J. Fitzpatrick, D. T. Jensen.




1-1

1.0 SUMMARY

The Apollo sp ace cra f t mission A-102 was successT'ully accomplished

in to ea r t h o rb it f rom com?lex 37B of

'3n September 18, 1964. The unmanned boilerplate spacecraft (BP-13) was

launched a t 11:22:43 a . m . e . s . t .the Eastern T e s t Range, Cape Kennedy, Florida, by the Saturn I Block I1

veh ic le , SA-7.

a

The purpose of t h e test was t o demons tr at e the cm ,ps t ib i l i ty of t h e

spac ecraf t with t he launch vehicle, t o determine th e lziunch and e x i tenvironmental parametc-rs f o r des ign ve r i f i ca t ion , and t o demons tr at e

the a l te rna te mode of escape- tower je t t i so n , u t i l iz in g the launch-escape

and pitch-control motors.

A l l mission t e s t ob jec t ives were f u l f i l l e d by t h e t im e of o r b i t a li n s e r t i o n , and addi t ional data were obtained by te lemetry th rough th eManned Space Fl ig ht Network u n t i l the end of e ff ec ti ve ba t te r y l i f e

dur ing the four th or b i ta l pass . Radar sk in t rack ing was continued bythe nztwork u n t i l the spacecra f t reen tered over the Ind ian Ocean d w i q

its 59 th o rb i t a l pass .

During th e countdown, th e re were no hol ds caused by th e s pacecrs t't.A l l spacecraf t subsystems f u l f i l l e d t h e i r s p e c i f i e d f u n c t i o n s t h r o u g h o d t

th e countdown and the planned f l ig ht - t es t per iod. Engineering da tawere received thrcnqh telemetry from a l l but two or' the instrumented

spacec ra f t measurements f o r the f u l l f l i gh t - t e s t pe r iod o f the miss ion .

The ac tuz l t r a j e c to ry a t time of S-I s t age cu to f f was s l i g h t l y 'higher than planned i n ve loc i ty , a l t i . tude , and f l ig h t - pa t h angle . A t

S-IV s t a g e c u t o f f , a l t i t u d e was s l ig h t ly lower and ve lo c i ty w a s s l i g h t l y

highe r than planned, which re su lt ed i n a more e l l i p t i c a l o r b i t t h a npi.anne$i.

The instrumentation subsystem was success fu l i n de te rmin ing the

launch and e x i t environment, and telemetry reception of t h e d a t a was

continuous through launch and ex i t except fo r a short Deriod during

veh ic le s t ag ing .

The launch-h?ating environment of th e BP-15 s pa ce cr af t w a s s i m i l a r

t o t ha t encounter zd by the BP-13 spacecraf t .

for th e two f l i g h t s were approximately equ al; however, th e inf lu ence ofs u rf ac e i r r e g u l a r i t i e s , as w e l l as c i r cumferen t i a l va r i a t ions i n hea t ing ,

Unless specif ied otherwise a l l t imes shuwn i n t h i s r epo r t are taken

Peak values a t most poi nts

a

from the ins ta n t of veh ic le l i f t -o f f ( launch vehic le ITJ umbi l ica l d is-

connect a t 11:22: 43 a . m . e . s . t . ).



1-2

was somewhat different for the two f l ig h t s because of d i f f e r e n c e s i nt r a j ec to r y and ang le of a t t a c k .

r a t e s were wi th in the predicted range. The heat-protection equipmenton the launch-escape subsystem (LES) was subjec ted t o temperatures

much lmr than the design l i m i t s which were es tab l i shed on the basisof an abort ed mission.

Both command and s e rv ice module h ea t ing

The launch-escape-tower jettison by the a l t er na te mode was success-

Pos i t iv e ign i t ion of the p i tch-contro l motor could no t be d e t e r -u l .

mined; however, the gene ra l t ra je c t or y ind ica ted t ha t it operated

properly.motor , c ar r ie d the tower s t r u c t u r e safe ly out of the path of the space-

c r a f t .

The launch-escape motor , toge the r with the pit ch-c ont rol

A l l str ain -ga ge, pre ssu re, and accel erome ter measurements indi ca te d

t h a t t h e spacec raf t per formed sa t i s f ac to r i ly i n the launch environment .Command-module c oni cal - su rfa ce s t a t i c p re ss ur es c or re l a t e d c l o s e l y with

wind-tunnel data , and the product of angle of attack and dynamic pressure(as) i d not exceed 1,000 (deg) ( lb /sq f t ) . The venting system of t hese rv ice m o d u l e performed as exp ect ed. The command-module i n s t r m e n t a c i o n

compartment d if fe re nt ia l press ure reached a maximum o f 13.3 ps i , bu tven ted r ap id ly a f t e r launch-escape subsystem sep ara tio n.

A 1.8g, peak-to-peak, 10-cps vi br at io n was noted during holddown.

Other v ibra t ion modes were similar t o those exper ienced dur ing the BP-13spacec raf t f l ig h t . One of th e simulated reaction-control-subsystem quad

assembiies w a s instrumented f o r v i b r a t i o n on t h e BP-15 s p a c e c r a f t f l i g h t .

The measured vibration l eve l s were above t h e design l i m i t .

The s t r a i n measurements i n the c m a n d module and s er vic e module

i n d i c at e d t h a t a l l bending moments were with in the design l i m i t s .

Of' t h e 133 measurements transmitted by telemetry from the space-

raf t , 131 produced continuous da ta .

The ground-support equipment performed s a t i s f a c t o r i l y duri ng pre-

laun ch and countdown o pe ra t on s.



2-1

2 . 0 INTRODUCTION

Apollo mission A-102 was th e second f l i g h t of an A p ~ ~ l l opacec ra f t_ ' . ; f igura t ion wi th a Saturn launch vehic le . The unmanned flight-test

veh ic le cons i s t ed OT t h e b c i l e r p l a t e 15 (BP-15) psc - c r a f t and the

SA-7 Saturn I Block I1 launch ve hic le . The space ve hi cl e, shown i n

f igure 2 .0-1 , was launched fram Cmplex 37B of the Eastern Test Range,

Cape Kennedy, Florida, on September 18, 1964, a t 11:22:43 a . m . e . s . t .

The BP-15 s p a c e c r a t t w a s the second of two bo i l er p l a t e spacecr af t

(see ta b l e 2 .0-1) p lanned t o be used i n demonstra t ing the comp at ib i l i ty

or' t he Apol lo spacec ra f t con f igu ra t ion wi th the Sa turn I Block I1 launch

v e h i c l e i n a launch and ex i t envi ronment us ing t r a j ec to r i es similar t o

those expected f o r fut ur e Apollo-Saturn V o r b i t a l f l i g h t s w it h p r od uc ti on

s p a c e c r a f t . The f i r s t of t h i s s e r i e s , t h e BP-13 s p s c e c r a l t , w a s success-

fu l? y launched on May 28, 1964 ( r e f . 1).

The spacec ra f t f l ig h t con f igu ra t ion cons i s t ed of a prototype launcb-

escspe subsystem (LES), boilerplate command module (CM), b o i l e r p l a t eservice module (a),nd in se r t . and adap te r . Bo i l e rp la t e f l ig h t - t e s t

sDacecraf t are developmental veh ic le s which simu late produ ction space-

c r a f t o nl y i n e xt e r n a l s ize , shape, and mass c h a r a c t e r i s t i c s .

B o i l e r p l a t e f l i g h t - t e s t s p s c e c ra f t are equipped with instrumenta-t i m t o o b t a i n f l i g h t - t e s t d a t a f o r e n gi n ee r in g a n a l y s i s a nd e v a l ua t i on .

These data are used t o confirm o r modify t h e d e s i gn c r i t e r i a f o r t h e

p roduc t ion spacec ra f t .

The f l i g h t sequence of major evsn ts dur ing the f l i g h t of t he BP-13s p a c e cr a f t i n t o orbit i s g i v e n i n f i g u r e 2.0-2. S p a c e c r a f t s e p a r a t i o n

from the launch vehi c le was n o t planned f o r t h i s f l! '. gh t; t he re fo re , t h esecond stage ( S - I V ) and instrument unit (IU) of the launch vehi cle, t c -get ,her wi th th e a t tached spacecra f t (wi thout the LFS which was j e t t i -soned) , were i n s er t e d i n t o o r b i t as a single u n i t .

v i s io ns o r p lans fo r recove ry of t he spacec ra f t .There were no pro-

The f l igh t t e s t of the BP-15 spacecraf t inc luded the fo l lowingf e a t u r e s n o t i nc lu de d i n t h e f l i g h t t e s t of t h e BP-l3 s p c e c r a f t :

(1) The in s t a l l a t ion of inst rum ents on one of the simula ted

reaction-control-subsystem quadrants on t h e SM f o r launch and ex i tt emperatu re and v ib ra t io n s tud ie s .

( 2 ) The demonstrati on of th e al te rn at e mode ot' LES j e t t i s o n , u s i n g

t h e lamch-escape motor and p i tch-contro l motor . ins tead o f using the

Cower- jet t ison motor , as i n t,he normal mode of j e t t i s o n .



2 -2

The f i r s t - o r de r t e s t ob j e c t i ve s fo r t he ov e ra l l m is s ion a pp l i e d

t o t he l aunc h ve h i c l e on ly , w i th t he e xc e p t ion of ve r i f i c a t i on o f

c ompa t ib i l i t y of t h e spacec ra f t wi th the launch vehic le under pr e f l ig hta nd f l i gh t c ond i t i ons .

The t e s t ob j e c t i ve s which a pp l ie d on ly t o t he Apollo BP-15 space-

craft were as follows:

(1) Determine the launch and exi t environmental parameters t ov e r i f y d e s i g n c r i t e r i a

( 2 ) Demonstrate the alternate mode of spa c e c ra f t U S e t t i s o n

u t i l iz in g th e launch-escape motor and pi t ch- con tro l motor.

j e c t i v e s were s a t i s f a c t o r i l y f u l f i l l e d .These ob-

An ev al ua ti on of the f l i g h t data has been made, and th e r e s u l t sof the evaluation a re p r es en t ed in this r e p o r t .



Mission

A-001

A-002

A- 101

A-102


TABLE 2.0-1. APOLLO SPACECRAFT FLIGHT HISTORY

Spacecraft

BP-6

BP-12

BP-13

BP-15

Description

Fi r s t pad abor t

High q abor t

Nominal launch andexit environment

Nominal launch and

exit environment

Launch da te

11-7-63

5-13-54

5-28-64

9- 18-64

2-3

Launch s i t e

White SandsMissile Range,

JT. Mex.

White Sands

Missile Range,N . Mex.

Cape Kennedy,Florida

Cape Kennedy,

Flo r ida

1, li




2-4 UNCLASSf FEED

Figure 2.0-1 .- Saturn-Apolfo space vehicle for mission A-102 at l i f t -off .

WN CLASSIFIED



c -5

120

100

80In0,

E

m0

Y

3

---.-2 6 0

a

.aJ0

.-U-a

40

20

0

Config- g.e.t. time , sec

Event uration Planned Actual

1. Saturn S -l lift-off

2. Saturn S -1 shutdown

Inboard engines

Outboard engines

3. Separation

4. Saturn S-IV ignition

5. LET jettison

6. Saturn S-IV shutdown

a 0 0

b

140 .9 141 .3

146 .9 147 .4

C 147 .7 148 .2

d 149.4 149.9

e 159.7 160.2

f 6 1 9 . 3 6 2 1 . 1

Range, nautical miles

Figure 2.0-2.- Sequence of major events for Apollo mission A-102.



3 . 0 FLIGHT TRAJECTORIES

Radars used

FPS-16

Azusa and ns-16

aFPQ-6 and ~ps-16

FPQ-6

1

g.e . t . , min:sec

o:oo t o 0:31

0:31 t o 6:O5

6 :05 t o 8:22

8:22 t o 11:04

The t r a j e c t o r i e s r e f e r r e d t o as "planned" were prefl ight-calcula ted

nomina l t ra jec tor ies suppled by Ma rs ha ll Space F l i g h t Ce nter (MSFC), and

t h e t r a j e c t o r i e s y e f er r ed t o a s "ac tua l" were based on the Manned Spzce

Fl ig ht Network t rack ing data . For both the planned and ac tu al t r a j ec -

t o r i e s , the 1963 revised Patrick a tmosphere w as used below 25 n a u t i c a l

miles and the 1962 U.S. Standard Model Atmosphere was used above

25 n a u t i c a l miles, except that t h e measured atmosphere a t t h e t i m e of

l i f t - o f f was used f o r t h e a c t u a l t r a j e c t o r y up t o 18.6 a u t i c a l miles.The e a r t h model used was t he F i sc he r E l l i p so id . The ground t rack for

the f i r s t thr ee o rb i t a l passes of the Apollo miss ion A - 1 0 2 i s presented

i n f ig u re 3.0-1.

three o r b i t a l passes, p re se nt ed i n f i g u r e 3.0-2, shows t h a t t h e a c t u a l

p r a f i l e was c lose t o th e nominal .

The a l t i tude- longi tude prof i le f o r the launch and

A comparison of th e ac tu al and planned mission event t imes f o r t h e

launch phase i s g iv en i n t a b l e 3.0-1. It can be seen from the t ab let h a t t h e a c t u a l S - I c u t o f f sequence was approximately 0.7 second l a t e rthan planned, and the ac tu al S - I V cutoff w a s approximately 2 . 0 seconds

l a t e r t ha n p la nned .

The a c tu a l l aunch t r a j e c to r y shown in f i gu re 3.0-3 was based on the

The data from

real-time data output of th e Range Sa fe ty Impact Pr ed ic to r Computer

(IP-7094)which used ~ 3 - 1 6 ,Azusa, and t h e FTQ-6 adars .

t he se t r a c k ing f a c i l i t i e s were use d du r ing t h e t i m e p e r i o d s l i s t e d i n

the fo l lowing t a b l e .

The ac tua l l aunch t ra jec tory i s compared with the planned launch

It can be seen from t h i s f igure that t h er a j e c t o r y i n f igure 5.0-3 .



3-2

ac t ua l l aunch t r a j ec to ry d id provide the launch envi ronment planned

for t h i s m is si on . A t S-I s t age cu tof f , t he a c tu a l t r a j e c to ry param-

e t e r s were approximately 198 f e e t pe r second h igh i n i n e r t i a l ve loc i t y ,1"

6,400 e e t high i n a l t i t u de , and approximately - h i gh i n f l i g h t - p at h2

angle.angle drop below the planned a t approximately T+5 minutes, and remain

below un t i l j us t be fore cu tof f . A t S-IV c u t o f f , t h e v e l o c i t y w a s3.5 e e t per second gre a te r , the f l ig ht - pa t h angle w a s O.OO23" less,and t he a l t i t ude was 1,704 f e e t less t han p lanned , r e su l t i ng i n a more

e l l i p t i c a l o r bi t .lower and the apogee was approximately 4 n a u t i c a l miles higher than

planned.

During t he S-IV burn ing , t he i n e r t i a l ve lo c i t y and f l i gh t -p a th

The perigee was approximately 0.2 n a u t i c a l m i l e

The or b i t a l por t i on o f t he t r a j e c to ry i s shown i n f i gu re 3.0-4.The p lanned o rb i t a l t r a j e c to ry was obtained using the nominal ins er t i oncondi t ions , suppl ied by MSFC, and integrat ing forward f o r th ree o r b i t a l

p s s e s .t h e o r b i t a l position and v e loc i t y vec to r ob t ained during t h e f i r s t pass

over White Sands Missile Range (WSMR). This vec tor w a s determined fronthe Manned Space Flight Network tracking data using the Goddard computer.

The WSMR vec tor w a s i n t egra t ed backward a long t he f l i g h t t r a j ec to r y t o

o r b i t a l i n s e r t i o n ( d ef in e d as S-IV cutoff p lus 10 seconds) and forward

f o r t h ree or bi ta l passes . These in te gra ted va lues were i n good ag ree-ment wi th the pos i t ion and ve loc i ty vec tors de te rmined by the Goddard

computer f o r passes near Pr eto ri a , South Afric a; Carnarvon, Au str al i a ;and Point Argue llo , Ca l i fornia , dur ing the f i r s t p s s . The i n e r t i a l

v e l o c i t i e s r e p o r t ed b y t h e s e s t a t i o n s a gr ee d w i t h i n 5 .5 f t / s ec , and t hef l i gh t - pa th ang l e wi th in 0.015". Thus the va l idi ty of t h e i n t e g r a t e d

o r b i t a l p o r ti o n of t h e f l i g h t t r a j e c t o r y was es t ab l i shed . It can be

seen in f igure 3.0-4 h a t t h e a c t u al o r b i t a l f l i g h t t r a j e c t o r y w a s i n

very close agreement with the planned.

The a c t u a l o r b i t a l p w t i o n of t h e t r a j e c t o r y w a s derived from

A comparison of the a ct ua l and planned t r aj ec to r y parameters i sg iv e n i n t ab le 3.0-11.di t i ons and o rb i t a l pa rame te rs were i n good agreement wi th t hos e

planned.

or b i t a l conf igurat i on , cons i s t i ng of t he BP-15 spacec ra f t , t he i ns t ru -

ment unit , and the %turn S-IV s tage , was c a lc u la t ed t o b e 53 o r b i t a l

passes, based on l i f e t ime drag cha rac t e r i s t i c s ob t a ined from SA-6 data.

The ac tu a l ree nt r y 3f t he o r b i t a l conf igura t i on i n t o t he Ind i an Oceanwas rzpor t ed dur ing t he 59 th o r b i t a l pas s .

The t ab l e shows t ha t t he ac tua l i nse r t i on con-

Using the WSMFi ve lo c i t y vec to r , t he e s t ima ted l i f e t ime of the

A comparison of t h e t r a j e c t o r i e s of the BP-13 and BP-15 spacec ra f tindicates that SA-6 had a higher ve loc i ty through max

SA-7 a t ta i ne d the higher ve loc i ty s ho r t l y before S-I engin2 8 cut-off

f o r an overa l l fas ter t r a j e c t o r y t h a n SA-6.

q and tha t



3-3

A complete detailed analysis of the f l i g h t trajectory of the l aunch

vehicle is presented in reference 7 .



3-4

TABLE 3.0-1.- MISSION IWENT TIMES

on l i f t - o f f s i g n a l as determined by launch-vehic leTIT umbi l ica l d i sconnec t a t 11:22:43.26 a.m. e . s , t . 1

.J

Event

L i f t - o f f

T i l t arres t

IECO

OECO

Ul lage rocke t s i gn i t i on

Separa t ion of $1 a.nd S - I V

S I V i g n i t i o n

U l l a g e r o c k e t j e t t i s o n

Launch-escape-tower jett ison

s - I V cu tof f

Planned,sec

0

136.3

140.7

146.7

147.4

147.5

149.2

159 5

159 5

619. 1

Actual,

sec

0

136 3

141.3

147.4

148.1

148.2

149 9

160.2

160.2

621.1

Difference,

s ec

0

0

.6

. 7

.7

.7

.7

.7

-72.0



3-6

Conditio n Planned Actu al

TABLE 3.0-11. COMPARISON OF SA-7 PLANNED AND ACTUAL

TRAJECTORY PARAMETERS - Concluded

Difference

P e ri g ee a l t i t u d e , s t a t u t e miles . . . .

P e r ig e e a l t i t u d e , n a u t i c a l m il e s . . .Apogee a l t i tude , s t a tu te miles . . . .

Apogee a l t i t u d e , n a u t i c a l m i l e s . - .

Period, min . . . . . . . . . . . . . .

Inc l ina t ion ang le , deg . . . . . . . .

115.17

100.08

136.75

118.83

88.58

31.76

kximum conditions~~

A l t i t u d e , s t a t u t e miles . . . . . . . .

A l t i t u d e , n a u t i c a l miles . . . . . . .

Space - f ixed ve loc i ty , f t / s ec . . . . .

E a r t h - f i x e d v e l o c i t y , f t / s e c . . . . .

E x i t a c c e l e r a t i o n , g . . . . . . . . .

E x i t dynamic pre ssu re, lb/sq f t . . . .Life t ime, revolut ions . . . . . . . . .

136.75

118.83

25,615.04

24,287.75

5.80

720. o

53

114.85 - 0.32

140.82

122.37

88.64

31.75

3.54

0.06

- 0.01

140.82

122.37

25,623.54

24,296.21

5.88

749.6

59

4.07

3.54

8.50

8.46

0.08

29.6

6



3-7

f'P

E

U N C L A S S IF1ED



3-8

al.-E

C0.-YIn.-E



3-9

8d

8d

8&0

8 s.-E

J

8

8s0

Le'2m

3c=U

IEm

rV

ma-cu0(

tc

.-&aa

.-E0-HUa3ccL

.EU

ie1

.-e

al

I

E.-

Ea

iicn



3-10

4FCI

8

8

8

S $c.-E

8d

8B

13

c



3-11

8d

8Q

3 s.-E

s

84

8B



3-12

80

0.-E

Pg 200

0

12

P 8

5c

a -

0o:oo01:oo 02:oo 03:OO 04:OO

Time, min:sec

(d) Dynamic pressure and Ma ch number.

Figure 3 .O-3.- Continued.

-1- ' W H D € W



7

8l i

E.-

.-c

a

3-13



S

em

0-d-3

.-

H

VII.-Es

U



m N 4 0 4 N mW IA 0

I I I

c.-Ei:r

-I-

3-15

d

a

EB8

a

t

f.PU



4-

4.0 SPACECRAFT DESCRIPTION AND PERFON!@JCF

4.1 Spacecraft Description

The Apollo boilerplate 13 (BP-15) spacecraft was composed of fourmajor assemblies

(CM), the service module (SM), and the insert and zdapteih.assemblies were similar in external configuration to the producLion

Apollo spacecraft.are shown in figure 4.1-1.in the spacecraft is given in figures 4.1-2 and 4.1-3.tem for orientation sild motion is given in figure 4.8-7.

the launch-escape subsystem ( L E S ) , the corrsnand module

These m j o r

The major assemblies and exterior dimensions of eachThe reference axes system f o r locations with-

The MSC axes sys-

The launch-escape subsystem (LES) i.s shown in figL1re it.1-4. Thetruss-type tower structure was a welded titanium tutukr frame, and theexposed surfaces were covered with silica-filled Buna-N rubber for ther-m a l insulation.command module by an explosive bolt. A OLructural skirt was mounted be-tween the top of the tower structure and the launch-esiape motcr. Thebolt attachments at the interface between the tower and the skirt pro-vided LES alinement capability. Two sequencers which forwarded a firing

signal to the L9S pyrotechnics were attached to the underside of theskirt

Each of the four legs of the LE S was at,tached to the

The LES motors f o r pitch control, tower jettison, and launch escapewere live. However, there were no initiators instaiied in the jettison

motor, and the wiring circuit from the sequencers to this motor was pur-posely not completed E O as to simulate a jettison-motor failure. Thealternate mode of toh-rr jettison (by firing only the launch-escape and

pitch-control mctors) was used.

A conical section of welded Inconel sheet was rnomted to the for-ward end of the pitch-control-motor housing.

183 pounds of sheet lead ballast to provide the proper LE S mass charac-teristics. The ballast enclosure also provided the Iriterface plane f o r

mounting the &-ball assembly. The performance of tb t : LES is describedin sections 4.5 and 4.6.

The section contained

The command module was conical with a convex base snd rounded apex.

The apex consisted of a fiber-glass radome containing the VHF telemetryomniantenna. (See fig. 4.4-2.) The CM sides were semimonocoque alumi-num structures terminating in the forward and aft heat shields. The ex-terior was covered with cork for protection against aerodynamic heating.Section 4.10 presents a description of the cork insulation configuration.

The inner side walls and the top of the cabin were insulated with aquilted glass-fiber material. The major components of the subsystems



4-

were mounted on shelves and brackets loca ted a long por t ions of t h e

inne r w a l l as shown i n fig ur es 4.1-5 and 4.1-6.

A cyl indr ica l a luminum st ruc ture w a s welded t o th e forward bulk-

head of t he CM t o s i mu l at e t h e egress tunnel o f the product ion space-

c r a f t .

P r io r t o launch , the ha tch was b o l t e d t o t h e CM e x t e r i o r s t r u c t u r e a n d

sea led with epoxy.

A main hatc h of aluminum al lo y provided access t o th e cabin.

External p rotuberances of the production s pace craf t conf igura t ion ,

inc lud ing the a i r ven t, umb i l i ca l f a i r in g , s imu la ted SM reac t ion- -contro l -

subsystem ( RCS) quad assem blies, and two sc im it ar antennas, shown i n

f i g u r e 4.1-7,were simulated fo r a be t t e r d e f i n i t ion o f l aunch env iron-

ment parameters.

The CM a f t h e a t s h i e l d w a s similar i n s i z e and shape t o t h e opera-tional heat s h i e l d . It was composed of an inner and o u t e r layer of lam-

i n a t e d g l a s s f i b e r over an aluminum honeycomb core and was a t t a c h e d t ot h e CM by four ad jus tab le s t r u t s .

used because the a f t heat; shield w a s not exposed to the launch environ-

ment and no recovery of t h e s p a c e c r a f t was planned.

.

f a i r in g , service -modu le s t ruc tu re , and in s e r t , a l l of which were bolted

toge the r . The b o i l e r p l a t e a d a pt e r was b o l te d t o t h e i n s e r t . The SM

assembly and the inser t , shown i n f ig ur e 4.1-8, were of semimonocoque

aluminum construction. For f u r t h e r s t r u c t u r a l d e t a i l s , s e e s e c t i o n 4 . 7 .

N o pro to type ab la t ive material w a s

The bo il er pl at e service-module assembly con sis ted of t h e CM t o SM

A pneumatically act uat ed umbi l ica l assembly w a s loca ted approxi-mately 18 inches below the top of t h e SM a t 122" ( f ig . 4 .1-2) .

e l e c t r i c a l power, ground-support-equipment (GSE) s ign als , and coolan t

f l u i d en te r ed the spacec ra f t t hrough the umbi l i ca l p r i o r t o l aunch.

External

Four s imula ted RCS quad assemblies were at ta che d t o t he upper por-

t i o n of t h e SM e x t e r i o r , 90" a p a r t .

namic cha ra c t er is t i cs of the production un i ts , the s imula ted un i t s were

similar in si ze and shape and were arranged on the SM i n th e same loca-

t i o n as they would be found on t he p roduc t ion spac ec ra f t . The RCS quadassembly loca ted near the -Z a x i s w a s instrumented t o provide temperatureand vibration measurements, For f u r t h e r d e t a i l s s e e s e c t i o n 4.8 .

I n orde r to dup l i ca te the ae rody -

I n add it i on t o the transduce rs and as soc iat ed components and wir ing,

t h e SM and adap ter contained e l ec t r ic a l wire harnesses which in t er f ace d

with the launch-vehic le ins t rument un i t fo r th e &-bal l , tower- je t t i son

command, and GSE s i g n a l s ( f i g . 4.1-8).



. 4-3

The spacecraft weight when inserted into orbit was 17.231 pounds;the spacecraft weight at lift-off was 23,838 pounds.

ment from Downey, the total weight of BP-15 spacecraft was reduced by1,600 pounds, which made the BP-13 and BP-15 spacecraf\ ballast config-urations almost identical. This reduction was ach; ved by removing bal-last from the SK and adapter. The BP-15 spacecraft weight was greater

than that of BF-13 spacecraft by 208 pounds at orbit incertion and

295 pounds at lift-off. This weight variation was que Lo known struc-

tural changes and manufacturing variations.module weight differences between the BP-13 and BP-15 Spacecraft. Thefinal weight was acceptable to Marshall Space Flight Center (MSFC).Actual weights for the command module, service module, and SM insertand adapter were individually obtained at Downey. The mted SM, CM,

and adapter were weighed together at the Eastern Tert Range ( E T R ) . Theactual weight and the location of the longitudinal c e n t e r c7f gravity of

the launch-escape subsystem were obtained at tte NASA- Plerritt IslandLaunch Area (MILA).until launch, and the actual weight and center-of-gravity data were ad-

justed.The weights shown in this table incl-ude ballast weight; of 2,014 ounds

i n the CM, 245 pounds in the adapter, and 183 pounds in the L5S.

Priclr to ship-

Table 4.1-1 hows major

Weight changes to the spacecraft were then moni torcd

The resultant mass characteristics are shown in table 4.1-11.



4-4

TABLE 4 . 1 - 1 . - WEIGHT COMPARISON OF BP-13 and BP-15 SPACECRAFT

Cormnand module

Service module

SM i n se r t and adap te r

To ta l spacecra f t i n o r b i t

LES

Tota l spacecra f t a t launch

Weight, lb

BP-13 I BP-15

9,300

4,172

A 2 2

17,023

6,520

23,543

9,477

4,149

31605

17 ,231

6 , 6 0 1

23,838

Difference

177

-23

54

208

-7 '

295



mu

8HP;w

1L n

rlI

!$

1

H

417'

2

r l

d

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d r i d

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rl

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R



4-6 U N C L A S S I F I E D

A

760"

t I \

Q-ball

Pitch-controlmotor ( l i ve )

\ Je t t i son m o to r

( l i v e )

1Launch

escape

subsystem

399.7"Launch-escape

motor ( l i ve )

VH F antenna radome

Command

fa i r ing I I

Z-

F igure 4.1-1.- Apol lo BP-15 spacecraf t .




Both views looking af t

F i n i l l

3-2I

Quad I l l

4-7

+Z

00

Quad IV Quad I

-ZF i n I

Total space vehiclecomponent Ioca ti ons(MSC nd MSFC)

t

-y 2700-@-900 +Y

I180'

-Z

Spacecraft instrurnentat onlocations (MSC only)

Figure 4.1-2.- Y- an d Z-axes and angular coordinate system used for

designating locations w ithin the BP-15 spacecraft.




4-8

. I nM M

rl

cv

a3b

M M a 3


0 I

0

I

. !

II

I

I

-

-IX

-

0

X

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aX

-

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-

r lInM I

ocoO bN r l

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Ind

U N C L A S S IF1ED



U N C L A S S I F I E D 4-9

Q-bal Iassetnb I Ba l las t enc losure

/

Tower je t t ison

-----otor ( I ive)

Launch-escape

motor ( l iv e) \

Launch-escape

tower-

Tower explosiveb o l t s

Pi c h-cont oI/’ otor ( l ive)

7 Towersequencers

Structuralsk i r t

-Y+zx+Z

Figure 4.1-4.- Launch escape subsystem for BP-15 spacecraft.




4-10U N CLASS1F I ED

.Eu

rdr

4

I

arnL

cc

L

0

Q)

8

.--4

t

-0S

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U NCLASS IF I ED 4-11


mcLn

Lnrl

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L

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.-



4-12 U N C l A S S IF I ED

- R & D VHF omniantennaand radome

Forward

LE S tower leg we l l

Aft comparttnent equipmaccess door (4 places)

X

-Z

Figure 4.1-7.- Command module exte rio r of BP-15 spacecraft.




I4-13

C -band

antennas Longerons

- / C M o S Mconnectors

Un s ru mented~~~

GSE umbi l ica l

(cover removedto show

connectors)

Instrumented

simulated RCSquad assembly

(1 place)

simulated RCSquad assembly( 3 places)

SM t o

insert

ib Service module

Iconnectors ,

Ring frame

(30 places)

Longerons

(6 laces)

Insert

Adapter

Stringers

Adapter to

f 2 8 places)

IU connectorsI

-Y -Z A i r c ond ti on ingbarrier

F igure 4.1-8.- Cutaway view o f BP-15 spacecraftse rvi ce module, insert, and adapter.



4-14 U "CLASS I F I ED

4.2 Instrumentation

Description.- The instrumentation subsystem provided conditionedanalog signals which indicated the launch and exit environment of the

BP-15 spacecraft to the communications subsystem.tation was supplied to satisfj measurement requirements in the following

areas: acceleration, acoustics, vibration, pressure, temperature, heatflux, strain, frequency, voltage, current, and discrete events.

Specific instrumen-

The block diagram in figure 4.2-1shows both the instrumentation

subsystem and the communications subsystem. Table 4.2-1 presents asummary of the measurement requirements for the Apollo mission A-102.Table 4.2-11 lists the flight equipment used to obtain the requiredmeasurements, and table 8.1-1 s a detailed list of the individual meas-urements on board the spacecraft. The location of the major instrumen-tation components within the command module are shown in figures 4.1-5and 4.1-6. The locations of most of the transducers are shown in fig-

ures 4.2-2to 4.2-7.

Also included in table 8.1-1are the ground-support- quipment (GSE)measurements and the transducer outputs which monitored angle of attack,angle of sideslip, and dynamic ram pressure from the &-ball system fur-nished by the NASA-MSFC. The six Q-ball outputs, two radial vibration

measurements, and one acoustic microphone output were routed to thelaunch-vehicle instrumentation unit and conditioned for launch-vehicletelemetry.

All of the measurements monitored on the BP-15 spacecraft requiredsome type of conditioning before they reached the modulation section.The instrumentation subsystem included two signal conditioning boxes,one low-level (0 to 10 mv) and one high-level (0 to 5 v).

The heat flux and temperature low-level signals were conditionedin the temperature signal conditioning box.

to the low-level 90 x 1~ ommutator and were sequentially sampled at

1.25 times per second. The signal conditioning box a l s o provided 0,

5 mv, and 10 mv d-c reference voltages to the 90 x 12;: ommutator to

enable accurate data reduction of the low-level system.

These signals were routed1

1

Amplification of the low-level signals was provided in order togive an output PAM wavetrain varying between 0 and 5 v d-c.

The high-level signals requiring conditioning were fed to the main

conditioning box. This box distributed the high-level signals to the

90 x 10 commutator and to the modulation packages. It also supplied



4-15

the commutator with 0 and 5 v d-c reference voltages f o r data reduction

purposes.puts 10 times per second and provided a PAM wavetrain varying between0 and 5 v d-c.

The 90 X 10 commutator sequentially sampled each of its in-

Both the PAM wavetrains and the continuous signals were fed to themodulation section where they frequency modulated their respective volt-age controlled oscillators (VCO)three modulation packages designated as A, B, and C. Package A containedVCO for IRIG channels 6, 9 to 16, and channel E.contained VCO for IRIG channels 6 to 8 nd 10 to 18.

VCO in each modulation package were mixed into two separate video com-

posite signals by two mixer amplifiers.composite output to the GSE for test purposes, and the other supplied

a composite signal to its associated transmitter in the W package.

The modulation section consisted of

Packzges B and C eachThe outputs of the

One amplifier supplied a

Configuration changes from B P - 1 3 spacecraft.- Shielding, as shownin figure 4.2-8, as added to the cabl.ing between the low-level commu-tator and the low-level temperature conditioning box input to eliminatethe electromagnetic interference (EMI) experienced on the BP-13 space-craft.the CM hatch removed and was adjacent to the service at.ncture on the

launch pad.

This interference was present only when the BP-13 spacecraft hac!

Strain-gage ranges were changed from 1,000 and 4,OOC pin./in. usedon BP-13 spacecraft to 500 @n./in. because of the low magnitude of meaE-urements experienced during B P - 13 spacecraft flight.

Sixteen temperature measurements and two vibration masurementswere added to instrument the RCS quad. "he addition of the vibrationmeasurements necessitated the deletion of two fluctuating pressure meas-urements so that the high-response channels could be used f o r the vibra-tion measurements.

Performance.- The flight performance of the BP-15 spacecraft i n s t r u -mentation subsystem was satisfactory. Of the 133 measurements instru-mented on the El?-15 spacecraft, only 2 failed to provide continuous data.Calibration of the thermocouple system used in the RC S is to be verified.

Data were lost f o r approximately 4 seconds during launch-vehiclestaging due to flame attenuation from the S-I stage retrorockets. No

interruption of RF transmission due to launch-escape-motor flame attei-uation was noted. Evaluation of the telemetry data received during thefirst orbital pass of the spacecraft indicated that instrumentation was

operating properly.



4-16

.--

Temperature measurement SR5877T, loca ted on the + pi tch nozz le of

the instrumented RCS quad (fig. 4.8-9) f a i l e d t o show the expected r i s e

in temperature.bandwidth w a s in di ca te d dur ing countdown and launch. This inform ation

l e v e l seen p r i o r t o l au nc h ap pe ar ed t o be normal f o r ambient conditions.

During t he f l i g h t , t he read ing was i ncons i s t en t wi th t he ac t ua l env iron-

ment determined by other thermocouple measurements in the immediate area.This measurement w a s s a t i s f a c t o r i l y c a li b ra t ed e l e c t r i c a l l y p r i o r t o

launch, b ut t h is was n o t a pos i t i ve i nd i ca t i on t ha t t he en t i r e t he rmo-couple c i rcu i t ry w a s f l ightworthy since the thermocouple was not included

i n t h e c a l i b r a ti o n c i r c u i t . For a d d i t i o n a l d e t a i l s rega rd ing t h i s meas-urement see sect ions 4.8 and 9.2.

A cons tant output of approximately &-percent info rmati on

.

Heat-flux measurement SA0553R lo ca t ed on th e SM under t he +pi tchnozzle of the instrumented RCS quad (fig. 4.2-6) provided qu es t ionabledata. Response w a s normal when the unit w a s ca l ibra ted approximate ly2 minutes prior t o l aunch which i nd i ca t ed c i r cu i t con t i nu i t y . Analysis

of th e data i nd i ca t e d t he t r ansducer r esponded t o ex c i t a t i on a t main-stage ignition. There w a s n o f u r t h e r response f r o m t h i s c i r c u i t , w h i l e

the corresponding body-temperature measurement SA0563T increased nor-mal ly wi th aerodynamic he a t ing dur ing th e f l ig h t . The exac t t i m e o rmode of lnalfunction w a s not appa ren t f rom the f l i gh t da ta . For addi-

t i on a l d e t a i l s regarding th i s measurement see sec t i ons 4 .11 an d 9.3.



U N CLASS IF1 ED

2

23

11-17

LESCM

SM

TABLE 4.2-1. MOLL0 MISSION A-I 32 MEASUREMENT REQUIREMENTS SUMMARY ,Ieasurement

Acce l e ra t i on

Vibra t i on

Pressure

F1 uct uat ng

I pressure

Acoust ic

S t r a i n

H e a t flu

Temperature

Quant i ty I Location

1

3

2

2

a

CM

SM

Adapter

SM-RCS

1

1 SM

2 Adapter

11 SM

I SM

1

Adapter

SM

Adapter

'Adapter

LE S

_.-

Requirement 1Determine s tr u c tu ra l body bending m0C.e

r esponses under f l i g h t l oads.

Determine s t ructural vibrat ion mode

responses under f l i gh t l oads .

ietermine s t a t i c aerodynamic loadingon ou t s ide con ica l su r f ace .

Determine f l i g h t p res sure ven t ing

c h a r a c t e r b s t i c s .- I

Determine aerodynamic loading on ex-

ter ior sui . I"aces and acoust ical

environment.

Determine ex t e r i o r acou s t i c l ev e l

under f l i g h t environment .~

Dete rmine s t ruc tura l stresses and

f l i g h t l o a d s .

Determine aerodynamic heating r a t e s

on ex t e r io r su r f aces .

Determine calorimeter body tempera-

ture f o r accura t e da t a r educ t ionof h e a t f l u .-- .~ _______-aerod.ynamic heating.

Determine in te r i o r temperature withi r*

SM and crew compartment.

Ver i fy proper heat t ra ns fe r from RF

packages.

D e t e r m i n e RCS tem pe rat ure s dur in g

i'l g h t .

______

~ -- _________

aTransmit ted on l a unch-ve h i cl e t e eme t r y .




4-18

A

20u

I

r i

4

H

I

HI

(u

-3-


UNCLASS IF I ED



U N CLASS IF1ED

Component Vendor Model

TM m odu la t i on package k n d i x TAP-316

90 x 10 commutator Appl ied Ele ct r oni cs 340-23-5

F i f t h Dimension LDAl2N-43290 x 1~ ommutator

Main s ignalco nd it i on ing box Brown CH-150

4-19

Quanti ty

3

1

1

1

TABL,E 4.2-11. - F L I G I T EQUIPJIENT FOR BP-15 SPACECRAFT INSTRUI.%NTATIONSUBSYSTEM

A m p l i f i e r r a c kd - c a m p l i f i e r

NASA-EEC 1 - A 11

Engin eere d Magn etics 2000D-1 7

Vibrat ion sys tem:

Accelerometer

a - c a m p l i f i e r

P res s u re s y s t em :

Transducer

d-c a m l i f i e r

Tem pera t u re s i gna l7-ond i t i on i ng box I Microdot

Endevc o 28191 3Endevco 242M5A 3

3

Statham PA-288TC-15-350 1 3R x i n e e r e d M a gn e ti c s 2000D-1 13

Endevco 2633MlOA

P r e s s u r e t r a n s d u c e r

Accelerometer

S t r a i n g a ge

Res is tance thermometer

Vibrat ion sys tem:

Accelerometer

a - c a m p l i f i e r

Vibrat ion sys tem:

Accelerometer

a-c a m l i f i e r

Wiancko I ~ 2-3236 -1 11

Donner i 4310 7

Baldwin-Lima-Hamilton EBF-13D 48

Trans-Sonics 2 1 6 8 ~ 92168A-2 82168~-2-12 1

T 408-2A-2 2

Endevco

b d e v c o

Endevco

7-T h e n oc oupl e Co1-t in e nt al TC-6 9

Sens ing TC-6A 3

C al o r i m e t e r W-Cal C-1123-A-5-012 aH y - C a l 1 C-1123-A-25-0 12

Zone box Microdot 401-01.38-1 2 0

Rugge de Fo rr es t 3878-2 12

FiicroDhone Gulton 5OIU0401 1

Acoust ic system: NASA MSFC

1I :

FCT-601Vm i t t e r - fo l l ower

a-c amplifierG l e n n i t e

~~

Endevc oEndevco

Ehdevco

1

28191

2633Ml3

2242M5A

2I :

I I

I Q-b a l l a s sem b ly 1 Nor t ron i cs 1 r-16 I 1 I




4-20 U NCL ASS IF1ED

1

I

I IU N C L A S S I F I E D




1

2

Measurement Location

1 L A O O l l A X ~ 3 8 0 Y O Z 6

2 L A 0 0 1 2 A X ~ 3 8 0 Y 6 z o3 C AOOOlA X c 78 Y O z 2 14 C A 0 0 0 5 A X c 78 Y O 2 2 15 C A 0 0 0 7 A X c 78 Y O z 2 16 S A 0 0 0 3 A X ~ 8 6 6 Y O z 73

7 S A 0 0 0 4 A x ~ 8 6 6 Y O 2 73

3

4

5

Note:Measurement nuiiibers refer t o

l is t ing i n table 8.1-1.

2700 9Oo-Y +Y

Far s i d e0 Near side

Figure 4.2-2 .- Loca tions of Ii i iear acceleration transducers for B P -1 5 spacecraft.

U N CLASS IF I ED



4-22 U N C L A S S IF1ED

b-JMeasurement

SA2120SSA2121sAAO 198sAAO 195sAA0197SAAO 196s

I

Location

X ~ 9 4 0 . 4 6 2 . 2 5 'X ~ 9 4 0 . 4 7 7 . 2 5 'XA 736 Y O Z 76X ~ 7 3 6 Y 7 6 Z OX ~ 7 3 6 Y 7 6 Z OX ~ 7 3 6 Y O Z 76

1

270° 9 Oo

-Y +YNote:

Measurement numbers refer to

l is t ing in table 8.1-1.

F ar s i d e0 Near side

Figure 4.2-3.- Strain-gage locations on B P - 1 5 spacecraft.




U N C L A S S F1ED

Measurement

1 SA0162P

2 SA0165P

3 SA0166P

4 SA0168P5 SA0171P6 SA0170P7 SA0172P8 AA0173P9 AA0174P

10 SA0169P11 SA0167P

12 SA0164P13 SA0163Pa

14 SA2760Y

27 0' 9-Y +Y

Far s i d e

0 Near s ide

/13

,

-1 1

--lo

D

4-23

L o c a t i o n

329.25'

277.5'

215.3O

187.25O277.25O316.6'277.25O183'3'58.9'147.9'

58.9'24.1'

0

a M i c ro p h o n e

i jo te :

Measurement numbers re fe r to

l i s t i n g i n t a b le 8.1-1.

F i g u r e 4.2-4 .- F l u c t u a t in g -p re s s u re t ra n s du c e r I x a t i o n s o n BP-15 spacecra f t .




4-24 U NCLASS IF I ED

1\ szyoo

4

-Y= 270'- t

I

Measurement

1 CA0078P

2 CA0076P3 CA0073P

4 C A 0 0 7 1 P-Z=18 0 5 CA0075P

6 CA0079P7 CA0074P8 CA0077P

9 9 CA0072P

6b'1

6

Locat ion

XI- 20 3570

35703570

180

180'

357;

9 3O87087'

Note:Measurement numbers refer to

l is t ing in table 8.1- I.Fa r s i d e

0 Near side

Figure 4.2-5.- Locations of conical-surface pressure transducers

on the BP-15 spacecraft,

UNCLASS IF1ED




Measu re men t

1 CA0580RCA0651T

CAO 58 1CA0652T

2 CA0582RCA0653T

3 CA0584RCA0655TCA0583RCA0654T

4 CA0588RCAO 659T

5 CA0591RCA0662T

6 SA0550R1

2 SA0 560 T7 SA0551R

3 SA056 1 T8 SA0553R

SA0563T4

9 SA0598RSA0 669T

5

10 AA0594RAA0665T

11 SA0555RSA0565T

1 2 SA0554RSA0564T

11 13 SA05529SA0562T

14 CA0589RCA0660TCA059 0 R

CA0661-I.' 15 CA0585R

CA0656TCA0586R

'-.--8 - I_ -

0w CA0657T270' 90 16 CAG587R-Y + Y CAO 658T

Location

xc 74

xc 74

xc 74

xc 53

X c 52

Xc52

Xc27

Xs338

Xs 3 1 5

Xs305

>(A933

X ~ 7 7 0

Xs267

Xs267

Xs305

Xc42.65

X c 2 7

X c 52

X c 52

Xc52

Far s i d e

0 Near side

Note:Measurement numbers refer tolisting i n table 8.1-1.

3O

180'

319'

3O

180'

3 19

319'

183'

187.2'

187.2'

183'

183'

145'

160'

177'

3 O

180'

80

85 '

95O

Figure 4.2-6.- Locations of heat-flux calorimeter body temperature measurementon B P - 15 spacecraft.





Location

1 SA0087D2 AA0090D3 AA0089D4 SA0088D5 SA0086D6 CA0021D7 SA0092D8 SA0091D

1

7

8

I I

Y -53.9 z 47.7Y -15.5 Z 7 1Y O Z 72Y 68.3 Z 22.8Y 92.8 Z 5 8Y 40.4 z 37.3191191'

I 1

2700 9 Oo

-Y S Y Note:Measurement numbers refer t o

l i s t ing in tab le 8.1-1.

(1)a r s i d e0 Near side

Figure 4.2-7.- Locations of vib ration transducers on BP-15 spacecraft.




C t 4-27



4-28

4 .3 El e ct ri ca l Power and Sequent ial

Electrical subsystem. - The e l e c t r i c a l subsystem provided the powerand c i rcu i t ry f o r t h e communications, environmental co nt ro l, instrumen-

tation, and sequencing subsystems.

ci r cu i t ry and components i s shown i n f ig ur e 4.3-1, with a layout showing

the location of the components within the cormnand module i n f igure 4.3-2.

A block diagram of the e lec t r ica l

Spacecraft power was provided by a n ex te rn al ground-support equip-

ment (GSE) power supply u n t i l T-12 minutes, a t which t i m e the power

loads were t r a n s f e r r e d t o t h e i n t e r n a l s p a ce c ra f t b a t t e r i e s . The bat-t e r y power fo r the ins tru men tat ion , environmental co nt ro l, and communi-

ca t i on subsystems came from two si lver -zi nc m i n b at te ri es A and B,

which were rated a t 120 amp-hr a t a 12-amp discharge rate .

The main ba t te r ies successful ly m e t a l l power requirements f o r themission including the f i r s t o r b i t a l p s s and exceeded the planned bat-

t e r y l i f e . Main battery A, which had a cur ren t dr ai n of approximately

17 amps, supplied 7 hours and 38 minutes of useful power.

v e r i f i e d b y t h e r ec e p t io n of t r a n s m i t t e r A a t t h e H a w a i i Radar Stat ion

on the f i f t h o r b i t a l p ass.

This was

Main battery B suppl ied u sef ul power fo r 5 hours and 20 minutes.

The cur ren t d ra in on ba t t e ry B w a s approximtely 25 amps from T-12 m i n -

u t e s t o T+5.6 minutes.

t e rmina ted, dropping th e curr ent dra in on ba t t e r y B t o approximately

l'j' amps. The useful l i f e of b a t t e r y B w a s v e r i f i e d by t h e recep t i on

of t r ansmi t t e r C a t th e Pre tor ia , South Afr ica , range s ta t i on on th ef o u r t h o r b i t a l pass.

A t T+5.6 minutes the Z S f a n operat ion was

The el e c t r i c a l subsystem fo r the BP-15 spa cecr aft w a s i d e n t i c a l t o

t h a t flown on t h e BP-13 spacecraft, w i t h the except ion of t h e prelaunch

condi t ioning of th e main ba t t er ie s and of th e sequencer pyro and lo gi c

b a t t e r i e s .e l e c t r o l y t e p er c e l l a s compared with 133 cc used on t h e BP-13 spac e-

c r a f t . The lo g ic and pyro bat ter ies flown on the BP-15 spacecraft were

used on th e second disch arge cycle, whi le those f l o w n on the BP-13

spacecraft were used on t h e f i r s t . These changes i n prelaunch condi-

t ion ing were based on the resul ts of b a t t e r y performance improvementt e s t s comple ted s inc e the BP-13 spacec raf t f l ig ht .

The BP-15 main batteries were ac t i va t ed us ing 135 cc of

Sequ entia l subsystem.- The se qu en tia l components co ns is ted of amission sequencer having two independent circuits (A and B) for r e l i a -

b i l i t y and two tower sequencers. A s shown i n f i g f l e 4.3-3,each tower

sequencer was c o n t r o l l e d by one of the two mission sequencer c i rcui t s .

The sequent ial subsystem included no bu i l t - in t ime delay between

r e c e i p t o f t h e s ign a l and i n i t i a t i o n o f t h e pyrot echni c s i n t he ex-

p los ive bol ts , launch-escape motor, and pi tc h- co nt ro l motor.



U N CLASS1F I ED 4-29

Using GSE power, the mission-sequencer logic and pyro circuits

were armed at T-7 inutes.Each circuit had one pyro battery and one

logic battery, which were rated at 6 mp-hr at a 1-amp discharge rate.

After arming, the voltages on the logic and pyro batteries were between

33-27 and 33.7 volts. Telemetry data indicated that the logic and pyro

buses remained armed and the batteries functioned satisfactorily through

tower jettison as required.

Twelve seconds after separation of the S-I and S-IV stages, a

signal was sent from the S-IV instrument unit to the sequencer logic

circuitry. The signal closed the firing circuits, and ignited theLES explosive bolts, launch-escape motor, and pitch-control motor.

Tower-separation command was confirmed by telemetry data which indicatedrelay closures in circuits A and B. Physical separation of the tower

from the Spacecraft was verified by the termination of the electricalmeasurements on the tower. Optical data conrirmed ignition of thelaunch-escape motor and separation of the U S tower.

The sequential subsystem flown cn tile BP-15 spacecraft was identi-

cal to that used on the BP-13 spacecraft and performed satisfactorily.





Nominal currentPower control box usage, amps

r

11 .25

1-wower amp 5.00

I I

Telemetry Bttansmi tter

1.25

5.0

Instrument

Telemetry VCO

GSE launchcontrol

1

I I

GSE power >ransponder no.

bus A

Transponder no. 2

--L Inverter- CS pump

I

--Inverter ECS fan

Barostat Total bu? Bbattery- SE I Spacecraft- Total spacecraft current

Figure 4.3-1. - Electrical power subsystem for BP-15 spacecraft.


4.5

17

1.25

5.0

1 o

1.0

9 O

8.0

25.25

42.25



U N C L A S S I F lED

~

4-31

Command moduleA/-

Figure 4.3-2. - El ec tri ca l power subsystem components.for BP- 15 spacecraft.

box




4-32

GSE launchcontrolequipment


GS E safe and arm controto driver circuits

Instrument uni t

Ground poin t

Tower bolts1Launch escape motor &

1 pitch control motor

Pyro bus A 2To circuit A

Ci rcu i t B same as circuit A@ Jettison motor switch

A Safe and arm driver

@ Logic jett ison relay

Figure 4 . 3 -3 .- Launch escape sequencer subsystem for BP- 15 spacecraft.




U NCLASS IF I ED

4 .4 Communications

4-33

Description. - The communications subsystem of the BP-15 spacecraft;

consis ted of th ree VHF te lemetry t ransmit te r packages and two C-bandtransponders (f ig . 4.2-1).

Each of t h e tra ns mi tt er packages shown i n f ig ur e 4 .4 - 1 c o n s i s t e dof a d-c t o d-c conv erte r, a 2-wat t FMtransmit te r , a 10-watt RF ampli-

f i e r , a n RF bandpass f i l t e r , and a power-line f i l t e r . The R F c a r r i e r

fr eq ue nc ie s were 237.8 mc, 247.3 me, and 257.3 me, and were de si gn at eda s t e l eme t ry l inks A, B, and C, r e spec t ive ly .

of each t r ansmi t t e r w a s s e t fo r *l25 kc.

The RF c a r r i e r d e v ia t i on

The RF ou tpu t s of th e th r ee t r ans mi t t e r pcka ges were f ed in t o amultiplexer which combined the individual outputs into one RF composite

s ignal .omi-anten na lo ca te d beneath th e radome a t th e apex of th e command

module (fig. 4.4-2).

Th i s s igna l was f e d by way of a bandpass f i l t e r u n i t t o a VHF

The power f o r th e A and B t ra nsm it t e r packages w a s suppl ied by themain bus A of th e s pa ce cr af t power system, and th e power f o r package C

w e s supp lied by t h e main bus B of t he sp ac ec ra ft power system.

requ i r ed fo r nomina l t r ans mi t t e r pe r fo mn ce w a s 6.5 amp a t 28 f 4 v d-c

per un i t .

Power

Two redundant C-band transp onders, shown i n fi gu re s 4.2-1 and 4.4-1,were used fo r t rack i ng th e spacecraf t dur ing launch, e x i t , and t h e i n i -

t i a l or b i ta l phases of t he mission .

t ransmit ted th rough a power di vi de r which, i n tur n, fe d or rece ivedfrom two he l i c a l antennas ( f ig s . 4.4-3 an d 4.4-4), which were flush-mounted, 180° apart,on t h e sur f ace of the service module. Power f o r

each of th e t ransponders was supplied by t h e main bus B of t h e space-c r a f t power system.

Each transponder received and

The transponders were i n t e r roga ted wi th an RF s i g n a l c o n s i s t i ngof two 1-microsecond pulses spaced 3.5 microseconds apart. The trans-

ponder decoder (fig. 4.4-4) r ece ived t he s ign a l and t r igge r ed th e

transponder tra nsm itt ers . Each of th e two transponder tra ns mi tt er s

generated a 0.77 microsecond p ulse i n reply.

mitted a t a minimum peak power output of 500 w a t t s .

These pulses were trans-

Performance.- During t h e launch phase of the BP-15 s pa ce cr af t, th eth ree te lemetry t ran smi t te r systems performed s a t i s fa c t or i l y and pro-

vided good quali ty data.

during th e launch phase occurred fo r a period of 3 seconds on a l l l i n k s

a t t h e time of S-I s t ag ing (T-I-148.2 sec ) .

maintained a t Cape Kennedy u n t i l W57O seconds.

The only interruption of RF t ransmission

Telemetry reception wasA complete l i s t of

U N C L A S S F1ED



4-34 U NCLASS I F IED

t e l eme t ry acqu i s i t ion and lo ss -o f - s igna l times f o r a l l r ange s t a t ions

i s r ecorded i n t a b l e 7.3-1.

During th e laun ch phasg, th e measured temperature extremes i n the

transmitt ing system were 43i n t h e maxirpm allowal$e l i m i t of 150' F, and they compare favorably

wi th the 40t h e BP-13 spacec ra f t ( ref . 1).t h e measured temperat ure extremes were 49' F and 56 as compred

w i t h 89' F and 120' F measured on t h e f i r s t pass of the BP-13 spacec ra f t ,

i nd i ca t i ng improved performance of t h e equipment coo ling subsystem.

F and 50' F. These values were w e l l with-

F and 55 3' extremes measured durin g th e launch phase of

During th e f i r s t pags over Cape Kennedy,

F

The two C-band tran spond ers, ca rr ie d on t h e BP-15 s p a c e cr a f t f o r

t rack i ng , were in t er ro gat ed during launch by th e Pat r ick Air Force Baseradar.l a s t e d f o r a period of 2 seconds a t t he t ime of launch-vehicle staging.

The only l o s s of t ransponder si gn al during th e launch phase

Throughout th e launch and or bi t al phases of th e BP-15 spacec ra f t ,t h e performance of t h e C-band transponders was good.t ion f requency (PRF), measured by w a y of te l emetr y l i nk A, of the t w o

beacons w a s normal.

The pu lse repet i -

The last r ange s t a t ion t o repor t acquis i t ion of the t ransponders

w a s H a w a i i on t h e t h i r d o r b i t a l pass. The C-band radar coverage timesfrom acqu i s i t ion t o l o s s of s ig na l fo r each of the pa r t i c ip a t in g r ange

s t a t i o n s i s shown i n t a b l e 7.3-11.

In conclusion, the communications subsystem su cc es sf ull y f u l f i l l e d

th e sp ec if ie d mission requirements, and performance w a s not degraded byei ther anomal ies or m l f u n c t i o n s .




U NCLASS F ED




4-36 U NCLASS1F I ED

Potted TMomniantenna ,

CM forwardcompartmentcover

Figure 4.4-2. - Locat ion of telemetry omniantennaon command module of B P -1 5 spacecraft.




4-3"

C-band antennatransponder A

+Z

I

C-band antenna

C-band antenna

X s 341-

X s 317-

XS 294-

xs 200-

Figure 4.4-3.- B P -15 spacecraft C-band beacon antenna locations

on the service module.



4-38

+360 v


He lic al antenna He1 ca l antenna

Powerdivider

Loca loscil lator

r:!

Duplexer

c 1

Pre-selector Magnetron

I ?

Pulse

networkI * Mixe r forming Telem etry

bTriggerdetector

vIF '

amp1 fierModulator

360 v -V

+720 v -b unit

+ 1 2 v -u

- 1 2 v -h

+ 12 v +

+12 v

+ 12 v

- 1 2 v 4> A A4- -1 2 v

Video Triggeramp1 fi e r generator

A

28 vd-c

4 lDecoderT-lI:::+720v

Power

supply

L ine

If i l terI +12v

Figure 4.4-4.- C-band transponder block diagram for BP-15 spacecraft.

-N C L A S S I F I E D



4.5 Pyrotechnic Devices

4-33

Pyrotechnic devices were used on the BP-15 spacecraft to ignite t.ielaunch-escape and pitch-control motors and to release the launch-escape

tower from the attachment to the command module. The assignments of Liepyrotechnic devices are shown in table 4.5-1.

The spacecraft prototype igniter cartridges (YA 453-0014-0081) wereinstalled in the launch-escape and pitch-control motors as shown infigure 4.5-1.bility for ignition. Each cartridge contained the Apcllo standard hotbridgewire initiator (M E 453-0009-0004) esigned to ig:n;te the propel-lant within 10 milliseconds when a current of 3 . 5 amps is applied tothe bridgewire.

Redundant cartridges were employed to enhance the relia-

The ME433-0014-0081 rototype cartridge performed satisfactorily 3nthe BP-6 and BP-12 spzcecraft (refs. 2 and 3 ) , and photographic dataindicated satisfactory performance on the BP-15 spacecraft.

The same cartridge conffguration except Tor a different thread size( l l E 453-0014-0082) gnited the tower- ettison motors satisfactorily onthe BP-6, BP-12, and BP-13 spacecraft (refs. 1 to 3 ) .

The launch-escape tower was secured to the command m’3dule by an

explosive bolt assembly in each tower leg as shown in figure 4.5-2. Thebolt assemblies were of an interim configuration, pending completion of

the development of the production spacecraft dual-mode bolts. The boltswere identical to those which were used successfully on the BP-6, BP-12,

and BP-13 spacecraft with the exception that the bolt body threads onthe BP-15 configuration were precision rolled to reduce stress concen-tration at the thread root, instead of being machine cut.

E ach b o l t assembly contained an ME 111-OOOl-OOl~ artridge ( s e e

fig. 4.5-3) ith dual initiators which ignited the single propellantcharge. The propellant gas pressure, operating against the actuating

piston (area ratio, approximately 20 to 1) compressed the silicone plugs.m e silicone plugs, under high pressure, served as a hydraulic fluid,

which loaded and broke the bolt in tension. The initiators were not theApollo standard initiators, but had the same response-time characteristicsand were of similar configuration. The only exceptions were a slightly

higher bridgewire resistance (1.35 + 0.1 ohm) and a larger ',bread size

to accommodate the interim explosive bolt. The Apollo standard initiators(1.00 hm bridgewire resistance) are planned to be used with the produc-tion spacecraft dual-mode bolts.

The firing current was applied simultaneously to th e explosive bci.tsand igniter cartridges of the launch-escape and pitch-control motors.



4-40

The bolts rel ease d t h e tower, 'and th e rock et motors prope l led the tower

out of the pa th of t he spacec ra f t .

The primary purpose of th e pyrotechnic devices was t o i g n i t e t h e

rocket motors and t o re le as e th e launch-escape tower. Opti cal and telem-

e t r y da ta i nd i ca t ed t ha t s epa ra ti on occur red a t ~+160.2 seconds.

c



4-42

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caunch escap e

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Figure 4.5-2.- BP-15 launch esca pe tower exp losive bolt installation.c.



4-44

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4-4:

4.6 Launch-Escape Subsystem Propul.sion

The launch-escape, pitch-control, and towcr-jettison motors usedin the BP-15 spacecraft launch-escape subsystem wcrc- of' the same con-

figuration as the respective motors used in the qualification test pro-

gram for Apollo Block I and Block 11.

To demonstrate alternate mode of tower jettison, the launch-escapeand pitch-control motors were utilized, and the tower-jettison motor

contained no initiators.

During a normal Apollo mission, the alternate mode of tower jettison

is planned for use in the event that the normal moan, using only the tcwer-

jettison motor, fails.

Launch-escape motor.- The launch-escape motor is designed to pro-vide the propulsive force required to remove the command module from

the launch vehicle in the event the mission is aborted during the count-down or launch phase through approximately 35 seconds of Saturn V second-

stage burning (at an altitude of approximately 293,000 ft). In additicn,

the launch-escape motor may be used in the alternate mode of tower

jettison.

The location of the motor with respect to the command and servicemodule is shown in figure 4.1-1,and the location with respect to thelaunch escape subsystem is shown in figure 4.6-1.diagram is shown in figure 4.6-2.

A rnotor configuraticn

The launch-escape motor used a case-bonded s o l i d propellant ofpolysulfide fuel binder and ammonium perchlorate oxidizer cast into aneight-point, internal-burning, star configuration. The motor had f o u r

nozzles spaced 90" apart and canted 35" outward from ti:e longitudinalaxis of the motor. The f o u r nozzles had graphite throat inserts andfiber-glass henolic exit cones. A nominal thrust vector off-set of

2.75" from the motor centerline was provided by the use of one oversizeand one undersize nozzle in the pitch plane. The thrusz-vector offsetwas provided so that the negative pitch thrust vector passed more nearly

through the center of gravity of the launch escape vehicle. PolprethEneblowxt closures were glued into each nozzle throat to provide a sealedenvironment inside the motor during handling and s t o r a g e .

The motor was ignited by a head-end mounted igniter, bhich incnrpcj-rated redundant hot-bridgewire initiators. The igniter was mounted inthe forward end of the motor, concentric with the longitudinal axis. 7'heigniter propellant was of the same formulation as the propellant used in

the motor. The initiators were used to ignite boron-potass'lum nitrate

pellets, which, in turn, ignited the igniter propellant.



4-46

The predicted performance parameters for the motor are presented intable 4.6-1. A predicted thrust as a function of time is presented

in figure 4.6-3.

Pitch-control motor.- The pitch-control motor is designed to pro-vide a positive pitching moment to change the initial attitude of thecommand module in order to remove the command module from the launcharea during a pad abort and from the flight path of the launch vehicle

during a flight abort. For the alternate mode of tower jettison thepitch-control motor provides the pitching moment required to assure pro-

per clearance of the launch escape tower from the launch vehicle andcommand module.

thrust vector angle of the launch escape motor for approximately 0.6 sec-ond to provide greater lateral displacement of the LES away from the

spacecraft p r i o r to LES tumbling.

The positive pitch thrust counteracts the negative

The pitch-control rn3tor used the same p r o p e l l a n t formulation as

the launch-escape motor.internal-burning, star configuration.

The propellant was cast into a 14-point,

The motor had one nozzle containing a graphite throat insert housedA polyurethane blowout closure was gluedn a steel structural shell.

into the nozzle to provide a sealed environment inside the motor during

handling and storage.

The motor w a s ignited by a pellet-type igniter which was mountedin the head end of the motor, concentric with the longitudinal axis.Redundant hot-bridgewire pyrotechnic initiators were used to ignite

boron-potassium nitrate pellets, which ignited the motor propellant.

Location of the motor with respect to the command and servicemodules is sh o w in figure 4.1-1, nd the location with respect to thelaunch-escape subsystem is shown in figure 4.6-1.tion diagram is shown in figure 4.6-4.

A motor configura-

Predicted performance parameters are presented in table 4.6-1. Thepredicted thrust as a f'unction of time is presented in figure 4.6-5.

Tower-jettison motor.- The tower-jettison motor is designed toprovide the propulsive force for removing the launch-escape subsystem

from the flight vehicle for a normal mission after approximately 35 sec-

onds of Saturn V second-stage burning, and from the command module duringan aborted mission, when the abort occurs before approximately 35 seconds

of Saturn V second-stage burning.

The tower-jettison motor had no function in the secondary mode oftower jettison, and on EP-l5 spacecraft the motor was not connectedelectrically to the sequencing circuit.



4-47

The motor configuration for this flight was basically of thequalification design.

mately forty Ti-inch-diameter high-shear bolts on 2-inch centers around

the circumference at each end of the interstage adapter to reinforce

the spot-welded flanges. These bolts were added following an inter-stage failure during the tower-jettison motor static test (see

fig. 4.6-1).

The only change was the addition of approxi-

1

The location of the motor with respect to the comqand and servicemodules is show. in figure 4.1-1, nd the location with respect to thelaunch-escape subsystem is presented in figure 4.6-1. A motor config-uration diagram is shown in figure 4.6-6.

Flight performance.- No motor instrumentation was used for the

BP-15 spacecraft launch-escape subsystem; therefore, actual motor per-formance is unavailable.performance of the individual motors.

See figure 4.6-3 and 4.6-5 for predicted

The ignition signal for the launch-escape and pitch-control motorswas relayed from the sequencer at ~+160.2 econds.graphic coverage of the flight revealed that the first noticeable flame

from the launch-escape motor occurred at approximately ~+160.2 econds,Positive ignition of the pitch-control motor could not be determinedfrom the photographic coverage. However, the general trajectory andtumbling rate of the jettisoned portion of the LES indicated that theperformance of the pitch-control motor, as well as the launch-escapemotor, was satisfactory.

A review of photo-

Recovery of the LES was not attempted; therefore, no postflightanalysis was possible.



4-48

I P a r a m e t e r

tf -

Launch P i t c h

e s c ap e c o n t r o l

TABU 4.6-1. BP-15 SPACECRAFT LAUNCH-ESCAPE PROPULSION

1

I g n i t i o n 3 e l a y, td , s e c . . . . . . . . . . . . . 0.045 0.015

T h r u s t r i s e t i m e , t f , s e c . . . . . . . . . . . . 0.090 0.090

B u r n i n g t ime, tb , s e c . . . . . . . . . . . . . . 3.230 0.625,

A c t i o n time, ta , s e e . . . . . . . . . . . . . . 6.200 1.005

Maximum c h a m b e r p r e s s u r e , pmax, p s i a . . . . . . 1,480 1.635

p b , p s i a . . . . . . . . . . . . . . . . . . . 1,296 1,403

Maximum t h r u s t , Fmx, lb f . . . . . . . . . . . . 161,703 2 , 800

Total time, tt, s e e . . . . . . . . . . . . . . . 9.900 1.250

A v e ra g e c h am b er p r e s s u r e d u r i n g b u r n i p g t i m e ,

A ve ra ge t h r u s t d u r i n g t u r n i n g time, F b , l b f . . . 147,500 2 , 3 9 0

T o t a l i m p u l s e, It, b f - s e c . . . . . . . . . . . 615,300 1,770

+

SUBSYSTEM MOTOR PREDICTED PERFORMANCE'

[A t 70" F g r a i n c o n d i t i o n i n g t e m p er a t u re a nd vacuum a t m o s p h e r i c ~res sure ]

IB as ed upon h i s t o r i c a l s t a t i c t e s t d a t a , p r o p e l l a n t b u r n i n g rate,

a n d m o t o r g e o m e t r i c c o n f i g u r a t i o n .

vl a u n c h e s c a p e

f-// /- Time

0 p i t c h c o n tr o l

mo t o r m ot o r

I g n i t i o n t h r u s tFi

0 F i r i n g c u r r e n t a p p l i c a t i o n ti me

Timea l l m o t o r s

100 p s i a



4-:+9

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Ba llas t enclosurecovera l las t-\

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Launche scape motor4

Launch-escape motor7

thru st alinementf i t t ing

Launch-escape-, 11tower

/j

Pitch-control motorrsupport assembly

'1. Reinforcing bol ts

/

Power systems andinstrumentationwire harness

+z# -Z-Y ;

Figure 4.6-1 .- BP-15 spacecraft launch escape subsystem.



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4-55

4 .7 St ruc tu res

Summary.- Examination of a l l spacec ra f t s t r a i n gage, p r e ssu re , andecce le r a t ion da ta ind ica te s th a t the s pacec ra f t per fo rmed adequa te ly i nth e launc h environment. Maximum va lu es of c ~ q di d not exceed

1,000 (d eg )( lb /s q f t ) (minimum allowable c ~ q 5,800 ( d e g ) ( l b / s q f t ) ) .

S t a t i c pressures measured on the CM con ica l surfa ce were i n agreementwith BP-13 sp ace cra f t f l i g h t and wind-tunnel data. The spa cec raf t SM

venting system performed as expected.

ment d i f f er en t i a l p ressure d i f f er ed ra d i ca l l y from th a t o f BP-l3 space-

craf t and reached a maximum value of 13.3 p s i

a t t ower j e t t i son .

were measured i n t h e CM dur ing the hold-down per iod p r i or t o l i f t -o f f .

These osc i l la t ions had a frequency of approximately 1 0 cps and damped

ou t r ap id ly a t l i f t - o f f . Adapter s t ra in-gage data show low amplitude

o s c i l l a t o r y s t r a i n s d u ri ng t h e same t i m e period.

The CM instrumentation compart-

p

X-axis os c i l l a t or y acc ele ra t ion s of 1 .8g peak-to-peak '

( external ' in ternal )

Power s pe ct ra l analy sis of Y- and Z-axis ac ce le ra ti on s show pre-

dominant frequencies of 16 and 40 cps duri ng th e boos t phase. The

16 c p s o s c i l l a t i o n was al so observed on t h e CM X-axis acce lera t ion

records and w a s noted dur ing the f l i g h t of t he BP-13 sp ce c r a f t .

e r s t ra in-gage da ta do no t show th a t s i gn i f ica n t os c i l l a t or y bending

moments were produced i n th e ada pte r.

Adapt-

Service-module vi br at io n records showed th e expected response t o

The

fl uc tu at in g pressure. Amplitudes and frequency di st r i bu ti on were s i m i -

l a r t o those observed during the f l ig h t o f t he BP-13 spacecraf t .

SM s t r a in -gage data show low- level s t r a in r es u l t in g from the s h e l l modiresponse with a mzximwn overall RIS value of 33 p,in./in.

Strain-gage data from t h e SM and adap te r ind ica te tha t bo th stati:a n d dynamic loads were of an acceptab le mgni tude .

Accelera t ion data from instrumentation in a n RCS engine nozzle

show lev el s of a cc el er at io n which were above the the design values.

An examination of nozzle vi br at io n and SM v ib ra t ion r eco rds ind ica ted

no s t r u c t u r a l f a i l u r e of t h e RCS assembly.

Fluct uat in g pressures on th e SM and adapter sur fa ces were recorded

and , in genera l , v er i fy the design spe ci f ica t io ns .

S t r u c t u r a l d e s c r i p t i o n .- The s tr uc tu ra l assembly f or BP-15 space-

c r a f t c o n s i s t ed of t h e fol low in g components: launch- escap e subsystem

(US), command module (CM) , s e rv i c e module (SM), n s e r t , and adapter .

(See f ig . 4.1-1.



4-56

The launch-escape subsystem con si st ed of a motor package and tower

t r u s s s t r uc t u re , as shown i n fi gu re 4.7-1. The launch-escape tower

cons i s t ed of a ~ A L - ~ Vi t an iu m welded truss s t r uc t u re wi th fou r mainlo ng it ud in al members of 3.6-inch-diameter by 0.127-inch t ubin g, andconnecting members of 2.5-inch-diameter by 0.050-inch tu bing.

tow er t r u s s s t r u c t u r e was 118 inches long and approximately 36 by 36inches wide a t th e nozzle s k i r t a t tachment and 47 by 51 inches a t t h e

CM attachment. The int e rf ac e between the launch-escape motor and t h e

t r u s s s t r u c tu r e was a s k i r t which w a s a semimonocoque s tr uc tu re i n th eform of a t runcated cone.

sheet which was a t t a c h e d t o four t i tanium longerons.

a t t a c h e d t o t h e t ow er a t t h e upper ends of t h e lo ng it ud in al members by

f o u r b o l t s .

l o n g i t u d i n a l members and a t t a c h e d t o t h e CM by means of f ou r pyro tec hnic

bo lt s. The launch-escape subsystem st ru ct ur al configurati on w a s i d e n t i -

c a l t o that which i s planned t o be used on the productio n Apollo space-c r a f t .

The

The skin w a s 0.140-inch-thick 6 U - 4 ~ i t an ium

The sk i r t was

Separati on housings were bo lt ed t o th e lower ends of t h e

The boilerplate command module w a s a semimonocoque-type aluminum

st ru ct ur e which cons is te d of sk in , s t r i ng ers , longerons, and f rames.

The ou te r sk in w a s 5456 aluminum with a t h i ckness of 0.190 inch.

An in ne r compartment con taining instrume ntation, an e l e c t r i c a l

power system, and b a ll a s t req uired t o provide proper weight and cen ter -o f - g r a v i t y l o c a t i o n w a s provided in th e command module. The comrnand

module w a s a t t a c h e d t o th e SMlongerons by three tension t i e rods.

Compressive load s were ca rr ie d by s i x pads, th re e of which were a l socapable of ta ki ng sh ear Loads. The CM-SM i n t e r f ac e connect ions weresimi lar t o production connections.

connection.

See f i gur e 4 .7-2 fo r a t y p i c a l

The bo i l er p la te s t ru c t ur a l assembly shown in f ig ur e 4.1-8 cons i s t ed

of the service module, 141 inches long; t he SM i n s e r t , 52 inches long;

and th e adapter , 92 inches long; making a n ov era l l lengt h of 285 inches.

In add i t ion , a f a i r i n g 10.75 inches long was a t t ac h e d t o t h e t o p o f t h e

se rv ic e module. The ou ts id e diameter of t h e assembly was 154.0 inches.

The types of co ns tr uc ti on of th e thr e e components were similar.

The semimonocoque structure of t h e s e r v i c e module cons is te d of an

alwninwn skin which was r e i n f o rc e d w i t h r i n g frames and longerons.

in se r t and adap te r had s t r in ge r s i n add i t ion t o the Longerons.longe rons were r ive ted t o t he sk in and extended the en t i r e l eng th of

each se cti on. The longeron s were made up t o two s t e e l te es jo ine d by

an aluminum web and were of constan t dep th i n th e i ns er t and adapter .

The depth of th e longerons i n th e ser v ic e module var ied l i ne a r l y from

a maximum of 17.2 inches a t th e to p end, where i t m t e d w i th t h e command

module, t o 5.50 inches a t the bottom end. The skin of a l l three com-ponents w a s made of s of a constan t

The

The



4-57

thickness of 0.16 inch over the entire boilerplate.

forced with a total of 30 ring frames.

components were all m d e from 2024-T4 aluminum with the exception of tbebottom frame of the adapter, which was made of steel. The ring framesin the service module were evenly spaced at approximately 12-inch inter-

vals, whereas the ring frames in the insert and adapter were spaced at.varying intervals of approximately 7 to 10 inches. In addition to the

six heavy longerons, there were a total of 28 tee-shaped stringers m6.e

of 7075-T6 aluminum in the insert and adapter sections.

The skin was reiri-

The ring frames in the three

The service module-to-insert and insert-to-adapter interfaces were

bolted connections consisting of 24 bolts each (4 olts at each longeron).The interface connection between the adapter and instrument unit con-

sisted of 32 bolts, evenly spaced around the circumference.

The parachute compartment of the command module was vented to thespace between the instrumentation compartment and the aft heat shield

by a connecting tube which passed through the instrumentction compart-

ment. This space was vented to the service module through the clearance

in the tension ties holes. (See fig. 4.7-2.) The service module wasvented to the atmosphere by eight holes, 4.9 inches in diameter, in theskin of the adapter at XA733.

coated nylon cloth was installed at the interface of the adapter and

the launch-vehicle instrument unit (IU). Thus, the venting of theservice module and instrument unit were independent of each other.

An air-conditioning barrier m d e from a

Launch winds and preflight aq p edictions.- Preflight aq pre-

dictions were based on the planned trajectory and upper atmosphericwinds obtained from rawinsonde soundings made at Cape Kennedy, Florida,at T-14 hours.

shown in figure 4.7-3.v a r i a t i o n s from T-14 to T-0 w e r e small as shown in f i g u r e 4.7-3.

predicted angles of attack based on the preflight trajectory and theT-0 wind profile are compared in figure 4.7-4 o angles of attack from

a preliminary evaluation of &-ball data by EFC.

aq

As can be seen, did not exceed 1,000 (deg)(lb/sq ft) d>xring hisflidht.which occurs at mximum dynamic pressures, was 5,800 (deg)(lb/sq ft).

The launch (T-0) nd the T-14 hour wind profiles areThe wind velocities were fairly low and the

The

A comparison ofprofiles using these angles of attack is also shown in figure 4.7-4.

This was a comparatively low value since the minimum allowable,aq

Command module static pressures.- Static pressures were measuredthroughout the flight at nine points on the surface of the commandmodule. The locations of the static pressure instrumentation are shown

in figure 4.2-3.T+gO seconds are shown in figure 4.7-5.flight-data coefficients with wind-tunnel data (ref. 6) at T+7O seconds

(M = 1.55) i s shown in figure 4.7-6.

Static pressure coefficients from TI-20 seconds to

A typical comparison of the

The correlation between the



4-58-.

flight-measured data and wind-tunnel data is good, and consequentlypressure loads obtained by the integration of these distributions agreed .

well with the predicted pressure loads.

Spacecraft venting.- The pressure inside the compartment made up

by the service module, insert, and adapter, and the pressure in thecommand module instrumentation compartment were measured and transmitted

throughout the flight.

The pressure inside the compartment m d e up by the service module,insert, and adapter was measured in order to verify the adequacy of theventing scheme used. This compartment was the only part of the space-craft with planned venting.spaced 4.9-inch-diameter holes located 13 inches forward of the inter-

face between the adapter and the instrument unit (fig. 4.7-7).

purpose of the SM compartment venting is the following:

The venting consisted of eight equally

The

-('internal 'external)1) To maintain the differential pressures

w i t h i n t h e bursting and crushing limits of the structure.

(2)tit a minimum.

To keep axial force relief provided by SM internal pressure

( 3 ) To maintain the differential pressure across the air-

conditioning barrier at less than 0.7 psi.

As shown in figure 4.7-8, he flight-measured pressure inside theservice module, insert, and adapter comprtment was slightly lower than

that predicted, except for the period from approximately ~+67oT+80 seconds. Recorded SM internal pressure histories from the BP-13spacecraft and the BP-15 spacecraft show close agreement.

son of the SM, insert, and adapter compartment pressure with that in theinstrument unit, as shown in figure 4.7-9, erifies that the differentialpressure across the air-conditioning barrier did not exceed the 0.7-psilimit. The data verify that the.venting scheme used was adequate.

The compari-

The command-module instrumentation compartment pressure was notrequired for verification of the venting scheme for the compartment madeup by the service module, insert, and adapter. The command-module in-strumentation compartment was not deliberately vented nor was it delib-

erately sealed. Figure k.7-10 shows the pressure time history of this

commrtment for both BP-15 spacecraft and BP-13 spacecraft. It can beseek from this figure that these pressure histories differ quite radi-tally. A maximum differential pressure of 13.3 psi peXternal< pinternal(was recorded prior to tower jettison.spacecraft instrumentation Compartment was of mGch tighter constructionthan the BP-13 spacecraft. It can be seen in figure 4.7-10 hat at the

It is apparent that the BP-15I



4-59

approximate time of LES j e t t i s o n ( ~ + 1 6 0 . 2 e c ) t he r e was a mrked in -

c r ease i n t h e le a k ra t e of BP-15 sp ac ec ra ft command-module instrumen-

tation compartment.

The results of a ven t ing ana ly s i s ind ica te that the observed de-crea se i n pres sure could have res ul te d from venting through a l e a k a r e a

of approximately 0.75 square inch. This le ak are a could have been ob-

ta i ned by one tower bo l t c le ar i ng i t s hole completely (f ig . 4.7-11) o r

by the looseni ng of t he washers around each bo lt , a l lowing venti ng

through the clearance around th e 8-inch-diameter bo lt in t he 1- inch- ,

diam ete r hol e. However, a r igorous explanat ion of the incre ased le ak

ra te a t tower separation cannot be offere d wi th the av ai l ab l e in forma-

t ion .

7

It should be noted t h a t th e tower bo lt , attachment scheme, and

cabin const ruc t ion used in BP-15 spacecraf t are not of production space-

c ra f t con f igu ra t ion .

Quas i -s t eady f l ig h t loads .- S ince the SA-7 t r a j e c t o r y was approxi-

mately a z e r o - l i f t t r a j e c t o r y , t h e q u as i -s t ea d y l a t e r a l loads were low.

The &ximum a,q experienced by th e BP-15 sp ac ec ra ft w a s

990 (de g) ( lb / sq f t ) compared wi th 4,600 (deg ) ( lb / sq f t ) f o r t h e BP-13spacec ra ft . I n e r t i a l l oads a t the in ter fac e of the adapter and I U

have been calculated a t t h e maximum c ~ q l i g h t c o n d i t i o n

(oq = 990 (deg ) ( lb / sq f t ) a t T+75 sec) . Time h i s t or ie s o f veh ic lelong i tud ina l and La te r a l acce le r a t ions are presen ted in f igu re 4.7-12.Bending moments resulting from a i r l oads a t mximum aq were a l s o

determined. The ne t bending moment on th e adapter-IU, inc lud ing a i r

6and i n e r t i a l l o a ds , w a s 1 . 31 X 10 i n - l b f o r t h e maximum c ~ q l i g h tcondition. hximum n e t load was small, approximte ly 25 percent ofthat experienced by t h e BP-13 spacecraft.

The axial fo rces a t XA 722 (adapter - IU in ter face) w e r e very s i m i -

l a r f o r th e BP-13 and BP-15 sp ace cra ft. The BP-15 sp ac ec ra ft experi-enced s l i gh t l y lower axia l loa ds tha n th e BP-13 sp ace cra ft because of

smaller angles of a t t a c k and dynamic pres sure s.

a t XA 722 ( a da p te r -I T i n t e r f a c e ) f o r t h e BP-15 spacec ra f t i s compared

with that cal cul ate d from f l i g h t data f o r both th e BP-15 and BP-13

spacec ra f t i n f igure 4.7-13. The predicted axia l force fo r the BP-15spacecraf t i s lower than the ac tu al for ce, due mainly t o t h e lower q

of t h e p r e f l i g h t t r a j e c t o r y and t o t h e small d i f f e re n c e i n p r e d i c te d

and ac tu a l SM in te rna l p r essu res .

P r e d ic t e d a x i a l f o r c e

/



4-60

LES pitch rates.- LES jettison and the resulting L;ES motions weredetermined from engineering film taken at Cocoa Beach and Melbourne,Florida. The initial LES motion was forward and translational in thenegative Z-axis direction. During pitch-control-motor thrust decay,

the U S pitch rate began to increase.initiation of the LES jettison sequence, the pitch rate had built up to

a steady-state value of 673 deg/sec.rates were 793 deg/sec. Stress analysis, with loads produced by theobserved pitch rate, shows that the tumbling rate was sufficient tocause yielding in the U S allast support plate. The analysis showsthat the tumbling rate was not sufficient to cause separation of theU S components from the main U S assembly. The quality of the best

films available was not good enough to verify that no components ofthe LLFS separated from the assembly.

Between 3 and 4 seconds after

Predicted LES steady-state pitch

X-axis vibrations.- An accelerometer system designed to measure

X-axis acceleration was located in the CM at the position shown infigure 4.2-2.and for a frequency response from 0 to 30 cps.

The system was ranged for amplitudes from -2. Og to 410. g

The X-axis acceleration records, from the measurement identifiedas CAOOIA, show a 10-cps oscillation having a peak-to-peak amplitudeof about 1-86 hich damped out rapidly at lift-off.The 1.8g value occurred after S-I ngine ignition but before vehiclerelease. Data obtained from the adapter strain-gage system show maxi-mum oscillatory strains during hold-down with peak-to-peak values of80 pin./in.

The maximum oscillatory X-axis acceleration recorded after lift-off

was about O.5g peak-to-peak and occurred at maximum dynamic pressure(T4-73 see).is shown by power-spectral-density analysis to be concentrated at afrequency of approximately 16 cps (fig. 4.7-14). The 16-cps oscillationwas also observed in the Y- and Z-axis acceleration records from theBP-15 and BP-13 spacecraft. Further analysis will be required to deter-mine the nature of this oscillation.

The majority of the X-axis vibration energy after lift-off

Y- and Z-axis accelerations.- The Apollo BP-15 spacecraft was in-strumented with six accelerometers to measure accelerations along theY- and Z-axes of the vehicle. Biaxial measurements were provided in theforward extremity of the I;ES, in the command module, and in the service

module. The LES or tower accelerometers were ranged for S . O g , and

CM and SM accelerometers for %.5g.

Tower Y- and Z-axis acceleration measurements at the limit of theinstrument range, @.Og, occurred periodically from 41 to 70 secondsafter lift-off. Examination of oscillograph records and narrow-band



4-61

analyses of the tower-accelerometer data show that the majority of the

response energy vas concentrated at frequencies of approximtely 16 and

42 cps.

The c o m n d module Y- and Z-axis accelerations exhibited similarwaveforms at corresponding times.erations show energy at 16 and 43 cps which indicates excitation offree-free body bending modes having frequencies of 16 and 43 cps. Ex-

amination of the power-specti-al-density plots presented in figure 4.7-15showed the majority of energy to have been concentrated at frequencieswell above that of the first few body-bending mod-es (first bending modefrequency z 1.9 cps). The measured response would not be expected tocontribute significantly to the overall structural loads, and no evi-dence of any significant oscillatory bending moment is shown on the

oscillograph records of adapter strain or RIG histories of adapter

strain.

Both comnd-module and tower accel-

As during the flight of the BP-17; spacecraft, the only excitation

of the first free-free vehicle lateral body-bending mode, observed onthe oscillograph records, occurred at S-IV engine ignition. Maximumtower-acceleration values during staging were below 0.8g peak-to-peak.

Estimated inertial loads produced by this oscillation would produceless than 5 percent of the design allowable bending moment (based onadapter-IU interface allowable loads).of the command and service modules verified all conclusions drawn fromthe tower Y- and Z-axis acceleration measurements since they exhibited

the same frequencies at smaller amplitudes.

The Y- and Z-axis accelerations

SM radial vibrations.- The service module was instrumented withthree accelerometers with a range of fsOg.

range of the accelerometer identified as SAOO86D was f r o m 20 to 1,000 cps.

Accelerometers identified as ~ 0 0 8 7 ~nd ~ ~ 0 0 8 8 ~ad frequency-responseranges from 20 to 790 cps. Two s imilar instruments were installed in

the SM adapter. A l l radial vibration accelerometers were installed onthe flar,ges of the frames adjacent to the skin to measure radial vibra-tion of the SM and adapter shells at the positions shown in figure 4.7-16.

The frequency response

Inspection of the SM radial vibration records showed random re-sponse as during flight of the BP-13 spacecraft.response, R E values of 11 to l5g, were determined from lift-off data.These vibrations died out rapidly as the vehicle rose, and the noise

environment became l e s s severe. The vibration response began to in-crease at "-I-40 seconds and continued to build up as .free-stream ynamicpressure increased and the vehicle approached transonic Mach numbers.Root mean square (RW) accelerations from SA0086D reached a mximum

of 21gT+55 seconds.

Random vibration

at T+5O seconds and decreased sharply to 7.5g R E atFrom T+55 seconds to T+74 seconds, the RIVE value remained



4 62

f a i r l y c o n s t a n t .

R E alue g ra du al ly decreased a s &ch number in cre ase d and dynamic

pressure decreased.

A f t e r maximum dynamic pressure a t T+73 seconds, t h e

The vibrat ion data recorded from SA0087D and SA0088D exhibited

the same general t r e n d as SA0086D with smaller RMS values.

a c ce l e r a t i o n values from ~ ~ 0 0 8 7 ~nd SA0088D were 18g and l5g, respec-

t i v e l y .

recorded th rough MSFC te lem etr y show t h e same gener al t re nd s as da ta

from SA0086D, SAOO87D, and SA0088D, bu t wit h s maller RMS v a lu e s i n t h e

comparable frequency range.

a r e p r e s en t ed i n f i g u r e 4.7-17.

Maximum RMS

A cursory examination of t he ad apter r ad i a l v i b r a t i o n data

RIVE h i s t o r i e s of t h e SM r ad i a l v i b r a t i o n

The rap id decrease i n rad ia l v ibra t io n ampl itudes no ted i n the

preceding paragraph may -Le a t t r i b u t e d t o t h e d ec re as e i n f l u c t u a t i n g

pressure which occurred a t the same time as the decrease i n v i b r a t i o namplitude (T+50 sec , M = 0.81) . T he decrease in fluctuating pressure

amplitudes was expected and can be assoc ia ted with changes i n t h e aero-dynamic flow a t t h i s Mach number and angl e of at ta ck .pressure and r ad i a l v ib ra t ion r eco rds from t h e BP-13 spacec ra f t show

t h e same t r e n d s as BP-13 spacec ra f t , and the t r end of f l u c t u a t i n g p re s-sure h is to ry i n t h i s Mach. number range has been v e ri f i e d i n wind-tunnel

t e s t s ( r e f . 6 ) .

Fluc tua t ing

The recorded v ib ra t i on response of t h e SM can be a t t r i b u t e d t o t h e

ex ci ta t i on of s h e l l modes of the SM i n se r t and adapte r s t ruc tu r e by th e

f l u c t u a t i n g p r e s s u r e s a c t i n g on t h e SM w a l l s . Evidence support ing this

conclusion i s presen ted i n t h e RMS h i s t o r i e s o f r a d i a l v i b r a t i o n and

f luc tua t in g p ressu res o f f igu res 4.7-17 and 4.7-18. A comparison of.these two plots shows very good t i m e co r re la t io n be tween t he RMS h i s t o r i e s

of r a d i a l v ib ra t i on and f l uc t ua t ing p ressu re.

Power sp ec t r a l ana l y s i s of vibrations recorded from SAOO86D and

SA0088D shows vibration energy concentrated a t 330 cps ( f ig . 4.7-19) .S p e c t r a l a n a l y s i s o f data from S A O O 8 p shows s imilar concentrat ions.

All r a d i a l vi br at io n datzL from th e BP-15 spacecraft show good agreement

with BP-13 s p a c e c r a f t d a t a ( f i g s . 4.7-19 nd 4.7-20). The narrow-band

analysis shown in f i g w e 4.7-19 shows energy concentrated a t resonant

frequencies which were determined i n ground vi br at io n t es t in g of th eBP-9 spacec ra f t ( a similar bo i l e rp l a t e veh ic le ) ve r i fy ing th e conc lusion

that the observed response was t h a t of s he l l mode-response.

The low level of SM s t r a i n da ta ( 3 3 pin. /in. R E ) recorded during

th i s f l ig h t ind ica te s th e observed she ll -mode exc i t a t ion d id not p ro-

duce stress levels which would damage th e BP-15 s p a c e c r a ft s t r u c t u r e .St ra in-gage power sp ec t r a l de ns i t ie s are -presented in f igure 4.7-21



4-63

and show a low level of strain energy.power spectral densities show energy concentrations at the same fre-

quencies.

The SM strain and acceleratior..

RCS engine nozzle vibration.- The BP-15 spacecraft was provided

with an instrumented simulated RCS quad assembly. A detailed descrip-

tion of the RCS assemblies is given in section 4.8.were mounted in the CW engine nozzle. The instruments were located as

shown in figure 4.8-5 with both accelerometers ranged at e0Og.

nozzle X-axis accelerometer SAOOglD had a frequency response range of

20 to 1,000 cps, and the accelerometer SA0092D had a frequency responserange of 20 to 450 cps.

Two accelerometers

The

The RMS histories of nozzle vibrations are presented in

figures 4.8-19 nd 4.8-20.

presented in figure 4.7-17 how excellent time correlation with theR.E histories of nozzle vibration which indicates that both SM andnozzle vibrations are excited by the same forcing functions.

design-load factors for these nozzles were 135g, 0-to-peak, at the

.center of gravity of the nozzle.lel to the vehicle X-axis and perpendicular to the nozzle centerline

in the vehicle Y-Z plane.

The RM3 histories of SM radial vibration

Dynamic

The load factors were applied paral-

Maximum acceleration value recorded from the nozzle X-axis measure-ment (SA0091D)was 174g, 0-to-peak, (T4-73.7 ec). Maximum accelerationvalue in the radial direction (SA0092D) was 160g, 0-to-peak, T+48 see).Correcting these maximum values to the design load condition (G levelsat nozzle center of gravity) yields a rnaximum of l5lg parallel to the

X-axis, and l3gg in the radial direction. The recorded values ofacceleration in both directions were above the RCS-mount-assembly designvalues. Resultant accelerations significantly higher than design values

can be determined by vectorial summation of the X-axis and radial nozzleac elerat on .

Powzr-spectral-density analysis of the nozzle X-axis and radialvibrations are presented in figures 4.8-21 nd 4.8-22 nd show themajority of the energy to be concentrated at frequencies which can beassociated with nozzle cantilever modes of vibration.

Vibration records from the nozzle and SM accelerGmeters do notgive any indication of a structural failure of the RCS assembly. Fur-

ther analysis is required to determine the significance of these vibra-tions.

Strain gages.- There were six strain gages mounted on the space-

craft structure. Two strain gages wer5 nounted in the SM and four inthe adapter. The SM strain gages were located to measure the circum-ferential tensile and compressive strains in the frames. The adapter



4 64

s t re in gages were loc a ted t o measure longi tud ina l t e ns i l e and compres-

sive s t r a i n i n t h e s t r i n g e r s .

The s t r a i n gages i n t he SM were l oca t ed on t he i nne r f l ange o f t he

AThe adept er s t r e i n gages were loca ted on th e i nne r f l ange o f

frame a t X

+Z-axis.

t he t ee - sec t ion s t r i nge rs on t he Y- and Z-axis a t XA 736 a t a rad ius

of 76 inches. Locati.ons of SM and adap t e r s t r a i n gages are shown i n

f i g u r e 4.7-16.t ab l e 8.1-1.

940.4 and c i rcumferent ia l ly a t 62.25' and 77.25' from the

Strain-gage ranges and frequency response a re shown i n

Data v a l i d i t y from t h e s t r a i n g ag es i s ques t ionable a f t e r ~ + 8 5 ec-

onds due t o the presence of thermal s t r ai n.

znalyzing the Emount of t hermal s t r a i n p re sen t i n t h e s t r u c t u r e i s avzi l -a31e.

No accurate method of

The SM st rain-gage system w a s ranged f o r k5OO pin./ in. A t i m e

h i s t o r y of t h e S M s t r a i n recorded by inst.rument numbers SA2121S and

SA2120S i s shown i n fi gu re 4.7-22. The s t r a i n g age i d e n t i f i e d as

SA2121S began t o measure dynamic strain a t approximately 0.1 second

a f t e r S - I i g n i t i o n w i th a peak-to-peak v alu e of 11 0 p,in./in. a t

T+O seconds.onds a t which time t h e s t r a i n s were to o low t o be read. Beginning a tT+41 seconds, t h e dynamic st r ai n increased t o a maximum value of

l3O pin./in. peak-to-peak, a t T+49 seconds, and then tap ered of f t o

zero a t approximately T+gO seconds.

s t r a i n data recorded from SA2121S a t a t i m e s l i c e from T+48 t o T+5O sec-onds showed that the m a x i m u m energy w a s concentrated around 330 cps,' as

shown i n f ig ur e 4.7-21. The osc i l l a to ry s t r a i n recorded f rom ins t rument

SA2120S showed ge ne ra ll y th e same t r ends as SA2121S excep t t he mgn i -

. t ude w a s somewhat smaller. .This can be seen i n the RlB t i m e h i s t o r i e s

o f t h e s t r a i n data rec or ded from t h e two gages shown i n f igure 4.7-23.

The s t r a in decreased from T+O un t i l approximately T+6 sec-

A power-spec t ra l ana lys i s o f t h e

The quasi-steady s t a t e l e v e l of t h e s t r a i n data from th e two SMs t ra in gages was determined by ut i l i z ing a 21-point averaging dataprocess ing rout ine ( see sec t i on 7 . 4 ) . From thi s ana lys i s , it was deter-

mined that the quas i -s teady s t a t e s t r a i n i n d i c a t e d b y SA2120S w a s s m l l

unt l l T+52 seconds a t which time the quasi-steady s t a t e s t r a i n i n c r e a s e d

t o approximate ly +4O win. / in. and then grad ual ly decreased t o zero a t

approximately T+73 seconds. Instrument SA2121S indica ted zero quas i -steady s t a t e s t r a i n unti l T+54 seconds, a t which t i m e i t decreased t o

approximately -20 pin. / in . and then gradua l ly increased to zero a tapproximately ~ + 6 7econds. This change i n quas i-stea dy s t a t e s t r a i ni s shown i n f igure 4.7-24. The var ia t ion of quas i -s teady s t a t e s t r a i ni n t h e se rv ice module frames may be a t t r i b u t e d t o t h e i n f lu e n ce of t h e

change i n t he ci rcumi ' eren ti a l s t a t i c p re s sure d i s t r i b u t i o n re su l t i ng

from th e change i n Mach number and a l t i t u d e wi th time.



4-65

The ada pte r strain-gag e system was ranged from *500 pin./in. The

t im e h i s t o r i e s of t h e s t r a i n measured by the four adapter gages are

shown i n f i g u r e 4.7-25.

t h e s t a t i c s t r a i n (due t o lg a x i a l l o ad ) of -18.4 pin. / i n . a t t h a t

loca t ion .

a t approximately T+67 seconds from instrument AA0197S.

s t r a i n data, the quasi - s teady s ta te f l ig h t loads , ax ia l f o r ce , andbending moment a t s t a t i o n XA 736 were determined. A t ime h i s to r y of

these loads a r e p r esen ted i n f igu re 4.7-26. In r e f e r r i ng t o these loads ,

it should be rea l i ze d t ha t t he accuracy of the loads de termined from

stra in-gage da ta i s questionable because of th e low mgn itu de of lo ads

and s t ra ins exper ienced dur ing the f l igh t .

The data g iven i n th i s f ig u re do not inc lude

The maximum compressive s t r a in measured was -120 pin. /in.From the adapter

Overall dynamic loads through t he launch phase of th e f l i g h t a tth e ada pte r s ec ti on were shown t o be very small, having a mximwn RM5

value of 5 pin. / in . During holddown, however, t h e a da pt er exp erie nce ds t r a i n s of 80 pin./in. peak-to-peak, with a m a x i m u m RMS value of

10 pin. /in.mined from a 21-poin t averag ing rou t ine) wi th rea l - t ime t o t a l s t r a i np l o t s w a s made. Th is comparison shows very small differences between

s t e a d y - s ta t e a n d t o t a l s t ra ins , ver i fy ing the conclus ion th a t no s ign i -

f ic a n t o s c i l la t o ry bending moments were produced i n the adapter .

A comparison of steady-state adapter s t r a in l e v e l s ( d e t e r -

Examination of a l l s t ra in-gage data i n d i c a te s t h a t t h e s t r u c t u r e

performed adequately i n th e f l i g h t environment encountered.

Fluc tuat i ng pressure . - The f luc tu a t ing pressures genera ted by th e

turbulen t f low over the spacecraf t SM were measured by means of 13 pres-

sure t ransducers .w i t h the frequency response of each shown i n t a b l e 8.1-1. The trans-

ducers were mounted s o as t o sense pressures on the SM e x t e r n a l surface.Locations of th e transducers are shown i n f i g u re 4.7-16.

The pressure range of a l l t ransducers w a s 0 t o 15 p s i a

Recorded pressures a t l i f t - o f f were i nva l id s ince the amb ien t

s t a t i c p re ss ur e w a s a t the upper l i m i t of th e transduce r range. A s

dynamic pressure increased during the f l ight , the re was a gradual i n -c rease in f l uc t ua t ing pressu re ampl itude un t i l W5O seconds, when asudden decrease occurred (fig. 4.7-18). The t r e n d of the data around

T+50 seconds i s ass oci ate d with changes i n t h e c h a r a c t e r of t h e l o c a l

flow a t t h i s free-stream k c h number (0 .9) and has been v er i f ie d by

wind-tunnel tes ts . After T+5O seconds, t he l ev el s ag ti in gradual ly in -

creased through maximum dynamic pressure (T+73 sec).

dynamic pressure decreased, the f luc tua t in g p ressu re l e ve l s decreased.

A s t h e free-stream

The highest n o i s e l eve l measured was i n t h e v i c i n i t y of an RCSnozzle.

di d not show th e drop of f a t T+5O seconds (f ig . 4 .7 -18(d) ) exh ib i t ed by

The f luc tua t i ng p ressu re h i s to ry in th e area around th is nozzle



.

the data'obtained rom all other transducers. This is attributed tothe unique locatlon of this transducer which was directly beneath a

forward-facingRCS nozzle as shown in figure 4.7-16 (measurement SAO165P).

R E I pressure histories from all fluctuating pressure transducersEre shown in figure 4.7-18.environment with Apollo design environment for several spacecraft zones

is given in figure 4.7-27.

A comparison of flight-measured noise



U N CLAS S IF1ED 4-67

-otor package

Structural ski r t

Tower truss structure

+z=+yY -2

Figure 4.7-1. - Apollo BP-15 spacecraft launchescape subsystem structure.

U N C LA SSI F I ED



4-68

Command module

Command module

-/-ervic e module

2longeron

Figure 4.7-2.- D eta i l of cornmand module-setvice module interfacefor BP-15 spacecraft.



UNCLASS1F I ED 4-69

uaJ

sUaJaJ

In

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n

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Y 'aPnt!llv




4-70

Yc.Q)-0aY

Y.--a

80

70

40

20

0 2 4

Angle of attack, af deg

7 on pref l ight nominal

trajectory and T-0atmosphere winds.

4 8 12x

aq (deg)( lb/ft2)

F igure 4.7-4.- Comparison of predicted to f l ight measured angles of attack and aq.1 0



4-71

.-I

dm5



4-72



4-73



4-74

1.6

1.2

.8

.4

0

a

1 .

- c

'1140 1100 1060 1 0 2 0XA, in .980

Figure 4.7-6 .- Command module static pressure coefficients for BP-15 spacecraft

flight data compared to wind tunnel data (reference 6) .



4-76

mu)

P

.-

c

Time, sec

Figure 4.7-8.- Service module internal pressure in B P - 1 5 spacecraft.1



4-78

uall

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a ir fl

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A Free atmosphere

4-79

Figure 4.7-11.- Possible air leakage path caused by fracture of explosive bolt .A



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4-84

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1 2 0

100

80

60

40

2c

0 2 0 40 60 8 0 I U

Time, sec

F igure 4.7-13.- Total axial force at interface of BP-15 spacecraft adapter andSaturn SA-7 instrument unit ( X ~ 7 2 2 ) .



4-85

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4-130

were e s s e n t i a l l y t h e same as f o r a pro to type i n j ec to r head.

valves and l ines were not : instal led for t h i s f l i g h t . The engine was

mounted t o th e q a d housing a t t he in j ec to r head s o t h a t t h e re w a s nodirect contact between the chamber and the housing.

pheno l i c - f ibe rg la ss lamina te in su la ti ng washer reduced thermal conduc-

t i on between th e i n j ec to r head and t h e housing.

shown schematically i n f i g u r e 4.8-5, i s i d e n t i c a l f o r p ro t ot y pe hardware.

P r o p e l l e n t

A 0.060-inch-thick

The method of mounting,

Quads B, C , and D were dummy ass emb lie s. The eng ine s co ns is te d of

s o l i d 4140 s t a i n l e s s s t e e l combustion chambers and throats which wereoverwrapped with a phen olic -f ib ergl ass laminate. The laminate extended

t o form t h e nozzle as shown i n f i gu re 4.8-6. These dummy engines wereheavier than the Quad A engines , bu t were s imi la r i n aerodynamic config-

u ra t ion and were o f t h e sane con f igu ra t ion as those f lown on the BP-13s p a c e c r a f t ,

4130 s t a i n l e s s s t e e l .

The quad hous ings were fabricated from 0,160-inch-thick

Instrumentation.- The RCS lns t rumenta t ion for th e BP-15 s p a c e c r a f tc o n s i s t e d of 16 temperature sensors mounted on t h e p o s i t i v e p i t c h ( + P) ,

CCW r o l l , a nd n e g a ti v e p i t c h ( -P) engines , and on th e housing s t ru c t ur e ,

a s shown i n f ig ur es 4.8-7 o 4.8-9. %o accelerometers w e r e mounted in

t h e CW roll engine nozzle 1:fig. 4.8-8). Tne +P, CCW, and -P engines had

tempera ture sensors lo ca ted on the nozzle , engine f lan ge , in je c t or head ,

and on the housing immediately below the engine. The +P and CCW enginesa l so have tempera ture sensors mounted i n t h e combustion chamber j u s t up-

stream o f t h e t h r o a t . A temperature sensor w a s mounted on th e unders ide

of the quad housing roof and on the engine-supporting bracket. The t e m -

pera ture sensors on the in , jec tor heads and on the engine suppor t ingb r a c k e t were resistance thermometers. The oth er temperature sens ors

Iwere Chromel-Alumel thermocouples contained i n a 8-inch-thick columbium

sheath ed ca se wi th in te rn al i ns ul at io n of magnesium oxide. The thermo-

couples were mechan ically clamped t o the chambers and housing.

was no d i r ec t con tac t of t h e Chromel-Alumel thermocopu le wires w i t h t h e

su r f ace o f the cnambers o r housings. The thermocouple mounting arr ang e-

ment and detai ls of the thermocouple design are shown i n f i gu r e 4.8-10.

There

The two accelerometers xere mounted in t h e CW engine nozzle perpen-

d i c u l a r t o t h e e ng in e a x i s , as shown i n f ig ur e 4.8-11.

measured both i n th e d i r ecz ion of spacec ra f t X-axis and i n th e d i r ec t ion

of an a x i s approximate ly perpendicu lar t o th e spac ecra f t X-axis. Epoxypotting compound w a s added t o secure t h e accelerometer cab les t o t h e CW

engine nozzle,

Vibra t ion w a s

RC S temperature.- in ce no RCS tem pe rat ure d a t a from t h e BP-13

spacec ra f t mission o r from wind-tunnel t e s t s were avai lab le for compar-

i son wi th FP-15 spacecraf t data, , a t h e o r e t i c a l a n a l y s i s w a s performed t o



pr ed ic t t h e maximum tempera ture s which would be a t t a in e d on the BP-15spacec ra f t RCS package. This an al ys is pre di ct ed maximum tempera tures

a t t he nozz le e x i t p l anes o f the +P, CCW and -P engines of 2,180' F,2,000' F, and 1,800' F, res:?ectively, which occurred a t T+154 seconds.

The predicted values a t the var ious sensor loca t ions are l i s t e d and

compared wi th measured valu es i n t a b l e 4.8-11. The large discrepancies,

evident between predicted a: id actual values i n d i c at e e i t h e r s i g n i f i c a n t

e r r o r s in th e ana ly t i ca l app roach used f o r the t empera tu re p red ic t ions

o r invalid temperature da ta .

Examination of t h e t empera tu re da ta ind ica te s th a t 13 of t h e 1 6The thermocoupleenp era tur e sensors functioned during th e mission.

l o c a t e d L inch from the exi-; plane of t h e +P engine nozzle (measurement

SR3877T) did not fun ction , and no data were ob ta ined fo r th a t locat ion.

The time-temperature da ta from t h e 15 functioning sensors are showni n f i g u re s 4.8-12 o 4.8-14,,rouped by engine.

from the engine-supporting 'tiracket and the underside of the quad-housing

roof are shown in d iv idu al ly i n f ig ure s 4.8-15 and 4.8-16.p l o t t e d as a ?-point average of tne indiv idu al data poin ts .

The thermocouple da t a

The d a t a are

A f t e r an i n i t i a l temperature drop res ul t i ng from aerodynamic cooling,

th e temperature of . t he +P engine began t o i ncr eas e a t T+40 seconds.

imum temperatures of 603' F and 745' F, as shown i n fi gu re 4.8-12, werereached a t t h e +P engine thi-oat and f lange, re sp ec ti ve ly, a t approximately

T+132 seconds. Maximum tempera tur es of approximately 1,080" F and 830' F,

occurring a t Ti-154 econds, were pred ic ted f o r t ne th roa t and f l ange ,

T h e +P engine i nj ec to r head, which w a s the h o t t e s t of t h e t h r e e i n j e c t o r

heads monitored, reached a maximum temperature of 159' F a t T+850 seconds.This i s well below the p red j c t ed value of 360' F.

Max-

The CCW engine temperature data , f i g u r e 4.8-13, a l so show an in i t i a ldrop, w i th t emp er a tu r e rise beginning a t T + 4 0 seconds, similar t o t h a t

o f t h e p o s i t i v e P engine. A maximum temperature of 775' F was reached

1 . 3 inches from t h e nozzle ex i t p lane of t h e CCW engine a t Tt-150 econds.This w a s th e highest temperature recorded by any of the sensors during

t h e f l i g h t , b ut i s considerab ly le ss than t he predic t.ed va lue of 1,865' F.

The CCW engine throat temperature reached a maximum value of amroximate ly

360" F a t T+160 seconds, as compared with a p r ed i ct e d w l u e o f 1,300' F

occurr ing a t T+154 seconds. The CCW engine i nj ec to r head reached E. maxi-

mum temperature of 142' F at, T+525 seconds and s tab i l ized at t h a t point ,

w e l l below the predicted value of 382' F.

The temperatures at ta ir e d on th e -P engine were f a r below predictel

values (see table 4.8-11), he highest temperature recorded for t h i s e n -

gine being 205' F on th e nozzle occurr ing a t T+200 seconds.

i n j e c t o r head reached a temperature of approximately 116" F a t T+850 se2-

onds and was s t i l l i nc reas i rg s lowly a t t h a t p o i n t .

The -P engine



4-132

The capabi l i ty of t h e quad housing and sup port ing bra cke t t o modu-

l a t e th e in je ct or head temperatures by permit t ing t h e h ea t t o d i f f u s e

throughout the s t ru c t ur e i s i l l u s t r a t e d by t h e r e l a t i v e i n s e n s i t i v i t y o fin je ct or head temperature t o engine f lange temperature. Although flan ge

temperatures varied from a maximum of 745' F on the +P engine t o a max-

i m u m of 122' F on the -P engine, th e thr ee i n je c to r head tempera turesremained within 50' F of one another through T+850 seconds.

t h e th ree in jec tor heads and t h e engine-supporting bracket a re shown i n

f i g u r e 4.8-17.CCW engine in jec tor heads had a t ta in ed an equi l ibr ium condi t ion in which

heat input from the chamber was equal to hea t output t o th e housing and

engine-support ing bracket .w a s s t i l l increasing a t T+850 seconds but w a s s t a r t i n g t o p eak a t t h a t

point .

manner a t T+830 seconds resul t ing from heat input from both the engine-support ing bracket and the -P chamber.and -P engine housings, and of tne quad-housing roof are spown i n f i g -

ure 4.8-18. The maximum temperature difference between these f o u rsensors w a s 55' F during t h e high-hea t ing per iod of launch; and a t

T+85O seconds, t h e four sensors were within 25' F of one another, in di-

c a t in g t h e c a p a b i l it y of t h e s t r u c t u r e t o d i s t r i b u t e t h e hea t input from

t h e chambers.

Data from

These da ta indica te that by ~ + 8 5 0econds, the +P and

The engine-supporting bracket temperature

!I!he -P engine i n j ec to r head w a s s t i l l i n c r e as i n g i n a l i n e a r

The temperatures of the +P, CCW,

Because of t h e discr epa nci es between pr ed ic te d and measured tempera-

tures , a su bs tan t ia l amount of addi t ion a l ana lys i s w i l l be required t oevaluate fu r the r t he i ns trumentat i on and t he a na l y t i ca l approach i n o rde r

t o re la te t h e data from t h e BP-15 spacec ra f t f l i gh t t o t empera tu res which

w i l l occur on l a t e r f l i g h t s w i t h prototype and qu al i f ie d hardware.r e s u l t s o f t h e a d d i t i o n a l a n al y s i s w i l l be given in a supplemental report .

Su bs t i t ut ion of prototyp e hardware i s expec ted t o re su l t i n i nc rea sed

nozzle temperatures since t h e t h inne r L-6@ nozzle w a l l of th e prototype

engine w i l l have a lower thermal capacitance and an increased thermalre si s t an ce t o heat conduct ion down and circ umf eren t ia l ly around thenozzle. The aluminum injector head temperatures w i l l a l s o be h igher onprototype engines s ince t h e 'thermal co nd uc tiv it y of t h e molybdenum com-

bustion chamber i s approximately nine t imes gr ea t e r than th a t of t he

L-605 combustion chamber of BP-15 s p ac e cr a ft , t hu s f a c i l i t a t i n g g r e a t e r

heat conduction in t o the in je ct o r head from th e chamber. The r e s u l t a n t

prototype i nj ec to r head temperature in crea se w i l l be tempered by the

add i t i on o f t h e p rope l l en t va lves t o t h e injector, which adds thermal

capaci tance t o th e inje cto r. The magni tude o f the se temperature in -c rea se s has not been establ ished a t t h i s t im e.

The

RCS vibra t i on . - The EMS t ime histories from the two accelerometers

mounted i n t h e CW engine nozzle are shown in figure 4.8-19 for the X-axis

and i n f i g u r e 4.8-20 fo r the perpendicula r axi s .

t o i nc re as e a t T+20 seconds, reaching 40g RMS a t T+44 seconds i n th eVibration response began



4-133

X-axis, and reaching

A sudden decrease in

51g REIS a t T+49 second s i n the perpendicula r axi s .

g l e v e l s w a s experi .enced during the period of T+49

t o T+54 seconds.of 47g i n the X-axis a t T+i’O seconds, approximately 45g in the perpen-

d i c u l a r a x i s a t ~ + 6 3nd T i 6 8 seconds.

follo wing maximum dynamic pr es su re , and ne it h er sen sor in di ca ted an out-

put beyond T+120 seconds.

design l i m i t .

between T+48 and T+50 seconds for tne X-axis and perpendicular ax is ar e

shown i n fi g u r e s 4.8-21 anc! 4.8-22. Most of t h e power i n bo th axe s was

concentrated around l9O cps, a t which frequency peaks of 100 t o 115 g /cps

occurred.which can be expected with prototype hardware.

i n th e r e p o r t r e f e r t o s e ct .i on 4.7.

The-RMS vibration level tnen increased t o a maximum

The vib rat ion le ve l decreased

The vibrat ion data obtained were above t n ePower-spect ral -dens i ty pl ot s for a 2-second time period

2

Further ana lys i s i s r e qu i re d t o r e l a t e t h e s e d a t a t o v a lu es

For a d d i t i o n a l a n a l y s i s



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2 77.2 5'

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-Z

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a

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Figure 4.8-1.- Location for service module RCS quads on BP-15 spacecraft.



Location of aluminumreinforcinq plate 7 I(under cork)

X s 2 9 4

cw

+ P engine

CCW engine

- CombustionS chamber

Throat

- ozzle

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187 2S0

Service module

Figure 4.8-2.- Service module RCS quad A on BP-15 spacecraft.

-Fm-



4-138



-Tapers from 0.044" at weld

to 0.030" at ti p -7

BP-15 RCS engine

1 nteg ra st ff c n ng r ngs 0.012" thick

4.400 Ti

L - 6 0 5 superalloy 1Prototype RCS engine

Figure 4.8-4.- Spacecraft BP-15 quad A and prototype RCS chamber design.

4-139



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X L Y to z Roll U P

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Z N x to Y Yaw $ W r

+X

t

Launch escape subsystem )I

Command module

+Z

36CS quads

Serv ice

y~ain hatch

moduleDirect ion of

launch azimuth

RCS quad A

' SpacecraftPosit ive maneuver & Linear Angular

Di ect ion A xi s Moment direct ion symbol velo city velo cityL I I

1 I I 1I

Figure 4.8-7.- MSC coordinate axes and notation system for Apo l l obo iler pla te and airframe, manned and unmanned spacecraft.

--u K Ass IF IED -



e---

+X

(Counterclockwise

SR7134T (Counterc lockwise

(Counterclockwise engine injector)

Nozzle

Combustion chamberbAUUY V J m o a t A

(Counterclockwise engine flange)

-P engine

(X-axis accelerometer)

S R 7 1 8 3 T

(Quad housing roo0

t

4 . 5 8 ” b Housing, 1

Clockwise engine

S A 0 0 9 2 D (Perpendicular axisaccelerometer)

I S R 5 0 6 5 T (E ngine supporting bracket)

Figure 4.8-8.- Servic e module RCS package instrumentation location sfor clockwise and counterclockwise r oll engines, engine supportingbracket, and quad hous ng roof on BP - 1 5 spacecraft.



4-144 45 clockwise engine

S R 7 1 2 5 T (-P ngine injector)

S R 7 1 2 2 T(-P engine flange)

I l l I \ I

L R 5 8 7 7 T

(+P engine nozzle)S R 7 1 7 5 T(-P ngine nozzle)

Combustionchamber SR587bT

(+P engine throat)I S R 7 1 2 1 T

Counterclockwise engine (+P engine flange)

S R 7 1 8 3 T

(Quad housing roof)

-P

S R 7 1 6 2 T ‘ S R 7 1 6 0 T

1 S R 5 0 6 S T

(-P ngine housing)

Engine supporting bracket

(+P engine housing?

(Engine supporting

bracket)

Figure 4 .8 -9 .- ervice module RCS package instrumentation locations for

+P and -P engines, engine supporting bracket, and quad housing roof on

BP-15 spacecraft.



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4 . 9 Acoustics

The Apollo BP-15 sp ac ec ra ft was ins t rumented t o measure t he ex te r i ora~311st icenvironment dur ing f l igh t . This environment con sis te d of

noise generated by the launch vehicle and noise generated aerodynamically

Available wind-tunnel da ta (r ef . 6 ) on the Apol lo spacecraf t conf igura-

t ion indicated that the aerodynamic noise environment could, for a nom-

i n a l Sa turn t r a j ec to r y , r each a maximum sound pr es su re l e v e l (SPL) of

16s dec ibe l s (d3 ) (rei': 0.0002 dynes/cm ) a t t h e CM-SM i n t e r f ace . Th i s

l o c a t i o n w a s instrumented on the BP-13 s p a c e c r a f t ( r e f . lo), and an

2SP L of 166 db ( r e f . 0.0002 dynes/cm ) was measured. It was a n t i c i p a t e d

t h a t t h i s s ound p r e s su r e l e v e l would a l s o be measured on the BP-15 space-c r a f t .

2

Theref ore, th e BP-15 s p a c e c r a f t w a s i ns trumented wi th t h i r t ee nf l u c t u a t i n g psessures and one microphone t o verify t h e wind-tunnel p r e -

d ic t ions . E leven of t h e f l u c t u a t i n g pressures and thz microphone werei n s t a l l e d o n t h e se rv ic e module and two f l uc tu at in g pressures on theadap te r . The t r ansduce r s are l i s t e d a nd t h e s p e c i f i c l o c a t i o n s arei n di c at e d i n f i g u r e 4.2-4 and table 8.1-1.

The p iezo ele c t r ic microphone sensed the app l ied d i f f er en t i a l

p ressu re , i . e . , pressu re va r i a t ions , and conver ted the p ressu re t o al i n e a r v o l t ag e t h a t v a s propor t iona l t o SPL.

The f l u c t u a t i n g p r e ss u r e t r a n s d uc e r s , however, were ab so lu te pres -

The pressures were convert ed t o a l i n e a r v o l ta g e.

sure gages . The ins t rument sensed the s t a t i c p ressure and tk s t a t i c

p r e ss u r e v a r i a t i o n s .Since th e pres sure tran sdu cer hsd t o accommodate atm ospheric press ure

a t l i f t - o f f , a n i ns tr um e nt w i t h a 0 t o 15 psia range w a s used. T h i s

range determined a I G w e r l i m i t f o r u seable f luc tu a t ing p ressu re da ta .

A B&K spectrum ana lyze r w a s used t o make overall and one-third

octave band SPL t i m e h i s t o r i e s for t h e f r e q u e n c y c a p a b i l i t i e s shown i n

t a b l e 8.1-1.da ta . By co r r e la t ing an o sc i l log raph r eco rd ing wi th the r educed da ta ,

t h i s d i sc r ep a nc y was reduced t o a s y s t e m a t i c e r r o r of 0 . 4 second.

A t i m e discrepancy was apparent i n t he reduced microphone

The overall SPL l e v e l time h i s t o r y for the microFhone i s presen ted

i n f i g u re 4.9-1.w a s i gn i t ed a t T-2 .4 seconds and reached a m a x i m u m o f 1 4 db

2( r e f . 0.0004 dynes/cm ) a t T+l second.

sound pressure l e v e l dec reased un t i l i t dropped below the ins t rumenta-

t i o n c a p a b i l i t i e s a t B l O seconds.

The launch veh ic le n ois e began when th e launch veh ic le

As t he veh ic le acce le ra ted , the

c



A one-third octave band an al ys is or' t h e launch vehi c le no ise a tW l econd i s presen ted i n f igu re 4.9-2. This analysis shows that the

maximum SP L had a broad band spectrum and i s r e l a t i v e l y f l a t f rom1-25 o 600 CPS.

Figure 4.9-3 i s a comparison of t h e overal l launch vehic le no ise

t i m e h i s t o r i e s for BP-15 and BP-13 spacec ra f t , and dep ic t s the s i m i -

l a r i t y between BP-15 and BP-13 launch vehic le SF'L.

difference between the two curves.

There i s very l i t t l e ,

The aerodynamic no is e measured on t h e BP-15 spacec ra f t was less

seve re than the no i se l eve l s measured on BP-13. The SPL exceeded the

ambient ins trum enta tion nois? a t W32.5 seconds ( f i g . 4.9-1) nd con-

t i nu e d t o i n c re a se t o a 149 3eci bels peak a t W43 seconds.

then decreased t c , 140 db a t T+47.5 seconds.

from N 4 7 . 5 t o W49.5 seconds; therefore, a one -thi rd octave banda n a l y s i s was made a t W 49.5 and i s presented i n f ig u re 4.9-4.rapid increase has been experienced during w i n d t u n ne l t e s t i n g of t h e

Apollo vehicle ( r e f . 6 ) and has been re l a t ed t o l o c a l sonlc f l o w con-

d i t i o n s a t the vehic le shoulder .

The SPL

An 18 db increase occurred

This

During supe rson ic flow, th e BP-15 spacec ra f t SPL remained a t approx-

ima te ly 15 0 db th rough the n a x i m u m dynamic pressure range.

seconds, th e aerodynamic SPI, dropped t o 140 dbt which ;,s considered i n -s i g n i f i c a n t .

A t W100

A r igorous explanat ion for th e lower aerodynamic l ev el s ind icat ed

has not been resolved . The vehi cle configu ratl ons were simi lar forBP-15 and BP-13 spacecraf t , , bu t the t ra jec tor ies p roduced a dir ' fe ren tangle of a t t a c k and dynamic pressure t ime history. I n add l t ion , th e

f l u c t u a t i n g pressures were relocated t o a i d i n s t r u c t u r a l analysis.Because of these d i f fer ence s , inc lud ing the ins t rumenta t ion lower l i m i tcap ab i l i t i e s desc r ibed above, add i t iona l inves t iga t ion and s tudy will

be required before a s o l u t i o n i s obtained.

F luc tua t ing p ressu re 3 a t XA 1000 and 329.25" nd f luc tua t ing

pressure a t X

and the h igh dynamic p r e s s r e range.

973 and 277.5" were ve ry ac t ive du r ing t r anson ic f l ig h tA

See section 4.7.

In summary, t he l aunch veh ic le no i se fo r BP-13 w a s s imilar t o t h a tfor B'P-13 spac ecra f t . Both reached a maximum of 14 8 db. (See f ig . 4.9-3.)The aerodynamic noise neasured on the BP-15 microphone resched a mxi-m um of 15 8 db ( f i g . 4.9-1) dhich i s l ess severe than the BP-13 microphone

measurement which vas 164 d b a t t h e same time i n f l i g h t . This d i f f e r -ence may be r e l a t e d to t h e 3 i f f e r e n t t r a j e c t o r i e s .




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4 . 1 0 Heat Protection

The heat protection on the BP-15 spacecraf t consisted of epoxy- '

impregnated cork covering the forward section of t h e command module and

t h e R CS housing and si l i c a - f i l l e d Buna-N rubber covering th e tr u s smembers of the la-mch-escape subsystem.

The forward section of the boilerplate command module was covered

wi th va ry ing th icknesses ( f ig . 4.10-1) f cork-based thermal insu la t i on .

This was t o prevent th e aluminum ski n of the command module from exceeding

th e desig n temperature of 250" F during the powered f l ight phase of t h e

mission.

X 133.72 because of t he ab l a t i ve qua l i t i e s of the f ib e r -g la ss radome.

The a f t hea t - sh ie ld area was not exposed t o the launch environment,and thus no heat p r o t e c t i o n w a s required.

No t he rma l in su la t ion was required between X c l 1 5 . 94 and

C

The command module (CM), service module (EN), i n s e r t and a d a p t e rwere not instrumented f o r s k i n temperatures .

Calorimeters mounted flu.sh with the exterior w a l l of' t h e CM and SM

were used t o measure launch-heating rates. A d e s c r i p t i o n of the launch-

he at in g environment i s covered i n sec t ion 4 .11 , Aerothermodynamics.

The he at pro te ct io n for the launch-escape tower con sis ted of Euna-Nrubber (60 p e r c e n t s i l i c a f i l l e d ) covering t he t r u s s members.p l i e s of rubber were b u i l t up eccen t r i c t o the s t r uc tu ra l tube wi th the

maximum thic kn es s i n t he regi on or' highest hea t ing . Truss members per-pendicu lar t o t h e flow were protected by a maximum thickness of

0.375 inch of r ubbe r .

maximum of 0 .3 inch of rubber .tower t russ members and t h e l o c a t i o n of th e temperature sen sors.

Sensors were mounted on.the metal sur fac es and covered wit h the Buna-Ninsu la t ion .

Several

Legs and diagonal members were pro tec ted by a

Figure 4.10-2 shows th e launch-escape

Figure 4.10-3 hows the measured temperatures a t t h e i n t e r f a c e ofthe rubber and the metal. su r f ace (bond l i n e ) during powered f l ig h t .The f igu re ind ica te s tha t a t t h e t i m e of tower j e t t i s o n t h e maximum

bond-line tempera ture on th e dia gona l member was 101" F.mented truss members which were p erpe ndic ular t o th e a i r flow i nd ica ted

a maximum temperature of 89' F. The bond-line temperatures of t h e mem-b e rs p a r a l l e l t o t h e a i r flow did not exceed 90" F. The temperature ofth e t ru s s member near est th e U S motor nozzle did not exceed 90' F a tthe time of tower j e t ti s o n . Temperatures measured during Apollo mission

A-101 (BP-13 spa cec raf t ) were wi t h in 10" F of t h e above temperatures.

Both instru-



4-180

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X c 2 4 . 1 2 -A f t heat sh ie ld

xco .o o -X c - 1 . 3

Figure 4.10-1. - Command module hea t prote ction forBP-15 pacecraft.



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Figure 4.10-2. - Launch-escape tower temperaturetransducer locations on BP-15 spacecraft.




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4 . 1 1 Aerothermodynamics

In order t o def ine the launch heat ing environment , 20 asymptoticca lo r ime te r s were ins ta l led on the spacecraf t : 12 on the command module,

7 on t h e s er vi ce module, and 1 on the adapter sec t ion ( f igs . 4 .11 -1 and

)+ . 11-2).

r a t e s b o t h i n c l e a n a r e a s and i n t h e v i c i n i t y o f v a r io u s s u rf a c e i r r e g -

Inst rumenta t ion loca t ions were se lec ted t o de te rmine hea t ing

;jlarities.

;I-,- command module and 0.9 B t u / f t /see on t h e service module are similart l T t h x e o b t ai n ed an t h e BP-13 s p a c e c r a f t ( r e f . I ) and ar e i n agreement

The maximum measured heating r a t e s o f 7.2 Btu/ft'/sec on2

-;th predicted VaLUeS ( r e f . 4).

A l l calor imeters appeared t o funct ion normally th rough l i f t -o f f ;

however, the calor imete r located on the ser vic e module behind th e fo r-

ward RCS nozzle d id no t respond t o the main heat pu lse a t 'rt-60 seconds.The ca lor ime ter i n the same posi t io n a l s ot h e BP-13 spacecraf t .

f a i l e d d u r i n g t h e f l i g h t of

spacec ra f t i s shown i n f ig -

the BP-13 spacecraft. The

diazeter exceeded 10 and had

"he f l i g h t environment f o r the BP-15

ure 4.11-3 and i s very similar t o t h a t of

Reynolds r x i b e r Sased on the raxinm, 5 d y4

dec reased t o 5 X 10 a t sta gin g. Hence, tu rb ul en t flow was expected

throughout that por t io n of th e t r a j ec to ry dur ing which heat ing occur red.

A Mach number of 9 .5 was reached a t t h e time of staging.heet ing r a t e s were g e n e r al l y a t t a i n e d a t a Mach number of 3.7 and a

The peak

/

bReynolds number of 4 X 10 .

The angle of attack began a gradua l ' i nc rease a t about W80 seconds

and continued t o increase th roughout the heat in g per iod ( f ig . 4.11-4).A more severe incsease occurred between T t l 3 O and W135 seconds. A l -

though Q-ball data during t h i s p or ti on of t h e f l i g h t are no t veryaccu ra te , t he data indicate the above trends.

of a t t a c k a r e also ind ica ted by t he ca lo r ime te r data .These t rends i n angle

Command module h eat inc . - The heating ra te h i s t o r i e s r ec or de d b y

the command module calorimeters are p re s en t ed i n f i g u r e 4.11-5 and aregrouped t o show c i r cumfe ren t i al va r i a t ions , va r i a t i ons a long con ica le lements , and the in f luence of var ious sur fac e i r r eg ul ar i t ie s .

A t a l l l oca t ions , the major heat pu lse w a s experlenced from w60to "-145 seconds wi th peak value s occ urr ing between W100 and W112

seconds.

t i on of ca lor imeters 7 and 12 t o 3 .3 Btu/f t2/sec a t t he loca t ion o f

Peak heating values varied from 7 .2 Btu/f t2/sec a t the loca-



4-186

calorimeter 2 , All ca lor imete rs except those a long the 180" c o n i c a l

element (2, 4, and 11) experienced a second peak a t approximately

T+135 seconds. The heating recorded by calorimeters 2, 4, and 11

dropped sharply as t h e h e at in g a t t h e l o c a t i o n s of t h e o t h e r c a l o r i -

meters approached t h i s second peak. A t T+135 seconds the Mch number

was about 7 and the Reynolds number was about 2 X Id.

An examinat ion of c i rc um fer ent ia l va ri at i on s and vari at io ns along

conica l el ements i nd i c a t e s an angl e -of -a t tack e f fe c t on l oc a l hea t i ng

which i s c o n s i s t e n t w i t h t h e t r ends o f t he &-ba l l data. T h i s e f f e c t i si l l u s t r a t e d i n f i g u r e 4 . l l - 5 ( h ) which shows t h e c i r c u m f e re n t i a l v a r i a -

t i o n i n h ea ti ng a t X,74. Calorimeters 1 and 3 located on the windward

s ide received the highest heat load, while calorimeter 2 on the leeward

s id e received th e lowest . This same effec t can be seen on ca lor imeters11 and 12 a t Xc27 ( f i g . k . l l - ? ( i ) ) and on calorimeters 4 , 5 , 6, and 9

a t X 54 ( f ig , k , l l - ? ( j ) ) .

T+100 seconds i s attributed to t h e inf luence of t h e s c i m i t a r antenna.The angle-of -at tack inc rea se began to influence heat ing a t approximately

T+gO seconds. The second peak on th e windward calor imete rs and th e cor-

responding rapid decrease on the leeward calorimeters (2 , 4, and 11)

occurred a t t h e same t ime (T+130 o T+135 seconds) as t h e r a p i d i n c r e a s ei n a n gl e o f a t t a c k ( f i g . 4 . 1 1 - 4 ) . I n the per iod f rom T+gO t o T+l23 sec-

onds, th e lo ca t io n of t h e windward c on ica l e lement va rie d between 360'and 345", and between 343' and 310' during t h e peri od from T+l25 seconds

t o T+135 seconds. This movement of t h e ef f e c ti v e wind vector caused th e

second peak a t th e loc a t io n of ca lor imeter 3 t o be higher than t h a t a tth e loca t ion of ca lor imeter 1.

The extended hea t ing a t calorimeter 5 a f t e rC

The magnitude of t h e heating a t these second peaks i s c o n s i s t e n t

w i t h p r e d i c t i o n s based on wind-tunnel data fo r an ang le of a t t a ck o f

5 ( re f . 5) .

Calorimeters 3 , 9, and 12 located downstream of th e tower leg wel l

i nd i ca t ed no s i g n i f i can t i n f luence of t h i s s u r fa c e i r r e g u l a r i t y on

lo ca l hea ti ng ( f i g . 4 .11-3(k) ) . No sharp dropoff i n hea t ing a t t h e

l o c a t i o n of calorimeter 3 , s imi la r t o that experienced by the BP-13

s p a c e c r a f t , w a s experienced during th e f l i g h t of BP-15 sp ace cra ft .

(See f ig . 4.11-7(a).) It appea rs t ha t a t t h e l o c a t i o n of c a l o r i m et e r 3,the f low w a s separa ted a f t e r approximately 95 seconds on th e BP-13 spa ce-craf t , whereas it w a s at tached throughout the heat ing period on the BP-15

spacecraf t .

Calortmeters 2 , 4, and 11 ( f i g . 4 . l l - 5 ( l ) ) i n d i c a t e no s i g n i f i c a n t

ef f e ct of th e hatch cover on l oc al heat ing; however, examinat ion of

th i s f i gure subs t an t i a t e s t he conc lus ions conce rn ing ang l e o f s t t a ck ,



4-187

The calor imeter nearest the apex on the leeward s ide (ca lor imeter 2 )

would be the f i r s t t o re sp on d t o an g le -o f -a t ta c k e f f e c t s as evidenced

by reduced h eat ing . Tt-mis e f f e c t c a n be noted dur ing the gradual in -c r e a s e i n angle of at ta ck from T+gO t o Wl25 seconds, as w e l l asdur ing the r ap id inc rease s h o r t l y t h e r e a f t e r .

A comparison of the data obtained a t calor imeters 5 and 10 on the

BP-13 and BP-15 s p a c e c r a f t ( f i g s . 4.11-6(b) and 4 . 1 1 - 6 ( ~ ) ) n d i c a t e s

th a t the upstream inf luence of the sc im i tar an tenna was grea te r du r ing

t h e f l i g h t of t h e BP-15 spacecraf t .

approximately the same on both vehicles; huwever, the locatr ion of

calor imeter 5 experienced a lower heat pulse on the BP-13 spacecraf t .

Dur ing the f l igh t of t h e BP-15 spacecraf t the heat ing a t ca lo r ime te r 5w a s approximate ly equa l t o th a t o f the hea t ing a t ca lo r ime te r 10.

(See f ig . 4.11-5(m). )

The heating a t l o c a t i o n 10 i s

The heating i n t h e v i c i n i t y o f t h e s t r a k e s t u b i s i l l u s t r a t e d i n

f igure 4.11-5(n) . As on the BP-13 spacecraf t , ca lor imeter 8 e g i s t e r e d

heat ing ra tes lower t han d id ca lo r ime te r s 6 and 7, which vere approxi-mately equa l. Fu r the r ana ly s i s i s r equ i r ed t o de te rmine t he r eason

for t h i s d i f fe r e nc e .

The heating caused by the LES motor during tower j c t t i s o n i s evident

a t a l l l oca t ions a t approximately ~ 1 6 0econds.

Service module and adapter heating. - The hea t ing - r a te h i s to r i e s fo r

the serv i ce module and adapter ca lor ime ters , p resen ted i n f igur e 4.11-7,

were s imilar t o those ob ta ined on the BP-13 spacec ra f t .compares th e hea tin g i n a "clean" area dur ing the two f l i gh ts .Figure 4.11-61:d)

All ser vic e module calor imet ers recorded decrease i n hea t ing a ta bou t T+135 seconds , a ga in r e f l ec t ing the e f f e c t of a sudden increasei n a ng le or' a t t a c k .

A s d ur in g th e f l i g h t of t h e BP-13 spacecraf t , ca lor imeter 13loc ate d on the surfa ce of th e ser vic e module behind th e forward RC S

nozzle did no t respond t o the main heat pulse a t T+60 econds. How-ever , t he ca lor imete r bodytemperature h i s t or y ind ica ted th a t th i s

area experienced h igher hea t ing than any o ther reg ion on the serv i cemodule, an obs erv ati on which i s consi s ten t wi th pred ic t i ons based on

wind-tunnel measurements (ref. 4 ) .

With the exception of' calor imeter 13, he se rv ic e module calorime2er

body temperatures remained within 6 percent of each o ther .

ature f o r ca lo r ime te r 13 was about 30 percent higher than the body tern-Derature fo r nea rby ca lo r ime te r s 18 and 20.

Body temper-

The maximum heat ing ra tes



4-188

for ca lor imeters 18 and 20 were about 30 percent h igher than the r a t e sf o r "clean" area calorimeters such as ca lor imeter 16.and 4.11-7(a) t o 4.11-7(d).

See figures 4.11-8,

Summary. - The launch heati ng environment of t he BP-15 spacec ra f t

was similar t o t h a t o f the BP-13 spac ecraf t . Peak valu es a t most lo -

ca t i ons were approxim ately equ al; however, th e infl uen ce o f surfacei r r e g u l a r i t i e s as wel l as c i rcumfe ren t i a l va r i a t i ons i n hea t i ng weresomewhat different during the tw o missions. Heating ra tes on both t he

command and s e r v i ce modules were within predic ted va lues .



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Heat flux (calorimeter) 9 C A 0 5 8 8 R . . 160.2 Tower jettison. ........

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(j) Calorimeters 4,5,6 and 9

Figure 4.11-5. - Continued.



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U N CLASS IF I ED 4-211

4.12 Equipment Cooling

Descr ip t ion .

-The equipment cooling subsystem for the BP-15 space-

c r a f t c o n s i s t e d of t h e equipment r e q ui r e d t o p r o t e c t c r i t i c a l e l e c t r i c a l

comDonents from over-heating. Thermal con t ro l or' these components w a s

provided by cabi n-air convective cooling and by coldp late conductive

cooling. There was no s imi la r i ty between t h i s subsystem and th e en-

vironmental cont rol subsystem f o r the Apollo production spacecraf t .

The equipment cooling subsystem was a pass ive , closed-loou heat-

tr a ns po rt system. The subsystem inc lude d five? co ldn la te s , a cab in -a i r

heat-exchanger and fan, a coolant st orag e tank, an accumulator , and aco ol an t pump. The e l e c t r o n i c equipment which w a s mounted on the cold-

p l a t e s c o n si s t ed of t h r ee t e l eme t ry FP ampl i f i e r and t ransmitter pack-

ages and two ra da r beacon tran spon ders . The cool ant fl u i d was a mixture

of 40-percent water and 60-percent e thy lene g lyco l ( spe ci f ica t io nMIL-E-3500). The water was d is t i l le d and deionized , and the mixtureconta ined no inh ib i to rs .

through a 35-micron abso lu te f i l t e r loca ted i n the ground suppor t equip-

ment. The subsystem sch ema tic i s shown i n fi gu re 4.12-1.

The water-glycol mixture was r e c i r c u l a t e d

During th e prelaunch a c t iv i t i e s , the equipment cooling subsystemwas ser vic ed with, and subs equentl y cool-ed by, the ground support equiF-

ment (GSE) model S14-052, wa ter- glyc ol coo li ng un it . The GSE c i r c u l a t e d

cold water-glyc ol through the spa cec raf t coolant stor age tank and main-

tained the stored coolant temperature below 20" F. During this sameper iod , the onboard coolant pump circ ul at ed t he cold wat er-g lyco l fromthe coolan t s to rage tank th rough the c o ldpl a tes and th e cab in heat ex-

changer co i ls . The warmed c ool ant th en r et ur ne d t o t h e pump and passedthrough th e thermal contro l va lve (TCV). When the coolant temperaturea t t h e pump o u t l e t was less t han 40" e" , t h e TCV rou ted the coolan t

d i re c t ly t o the c o l dd at es . When the pump ou t l e t coolan t tempera tureexceeded 40° f5" F, the TCV rou ted coo lan t f lu id by w a y or' t h e s to rage

t a n k t o t h e c o l d pl a t es . Thus, i n e f f e c t , a volume oy c o l d f l u i d was

withdrawn from t he s to r age tank t o rep lace th e volume of w a r m coolan t .

The temperature of t h e r e c i r c u l a t e d c o ol a nt t o t h e c o l d p l a t e s was main-

tai ned below 40" k5 " F, by the use of th e TCV, u n t i l t h e temperature of

the coolant i n the s torage tank reached 45" F, a t which t i m e , t h e TCV

was main tained i n the open pos i t ion t o allow a l l c a o la n t t o flow from

the pump through th e tank and t o the col dpl ate s.

Cooling of the water.-glycol by GSE ceased a t T-18 seconds whenthe spacecraf t umbi l ica l was disconnected.

parameters which measured the subsystem performance w a s a l so d i scon -

t inued a t t h i s time.and a f t e r imbi l ica l d isconnect .

Monitoring of some of t he

Table 4.12-1 shows the psrameters monitored before

'U N C LA SSI F I ED




The equipment cooling subsystem w a s designed t o main tain the cab in-

air temperature belaw 100" F thro ugh launch and t o mai nta in t he commun-

ic at io ns equipment below 150" F f o r one o rb i t a l pass , o rovided th e coolan ts t o r e d i n t he t a n k w a s below 20" F a t umbi l ica l d isconnect . These c r i -

t e r i a i n c l u d e d a 60-minute hold a f t e r umbi l i ca l d i sconnec t.

Af ter um5i l ica l d isconnect , the spac ecra f t sys tem cont inued t o

c i r cu la te the wa te r -g lyco l . A t 8 cabin pressure of 5.45 kO.5 p s i a ,

two baroswitches tu rned of f th e power t o the s ing le-phase inv er t er ,

which powered th e ca bi n fan. When th e fa n tu rn ed o f f , convective

cool ing of the cab in ceased .

The configuration of the equiment cool ing subsystem of t he BP-15spacec ra f t was s imilar t o t h a t of BP-13 spacec ra f t .uration change was that t h e BP-13 spac ecraf t cab in tempera ture probe

was r e l o c a t e d t o s e ns e t h e i n l e t g as -s tr ea m t e mp er at ur e i n t h e c a b inhea t exchanger r a t he r than the o u t l e t gas-stream temperature. The

probe was r e loca ted t o ena' cl e the sens ing of a more rep resen ta t ive gastemperature.

The only config-

Several hardware changes w e r e made t o t h e system as a r e s u l t of

the exper ience ob ta ined from t h e BP-13 spac ecraf t ground and f l i gh t

t e s t s and BP-15 spa cec raf t ground te s t s . The cabi n heat-exchanger

bypass valve on BP-15 spacec ra f t w a s permanent ly posi t ioned t o bypass50 percent o f the coolan t eround th e heat-exchanger core. On th e BF-13s p a c e c r a f t t h i s v a lv e o p e m te d t o d e l i v e r c ol d wa t er -g l yc o l t o t h e h e a texchanger core when the cabin temperature w a s grea te r than 70" F.

B ~ l m cabin-a i r tempera ture of' 70" F, 50 percent of the water-glycolwas bypassed around th e core. During th e BP-13 spacecraf t checkout,

the c ab in-a i r tempera ture taken a t the heat -exchanger ou t l e t , w a sfound t o be 45" t o 55" F.fo r su s ta ined ope ra t ion of t h e b a t t e r i e s .f o r th e cab in fa n and coo lant pump was manually programed during the

BP-15 spsc ecraf t launch countdown t o main ta in th e cab in a i r temperature

a t approximately 75" F and the col dpl ate -ou tle t temperature below100" F ( f i g . 4.12-2).

This temperature was judged t o be t oo co ld

A s a r e s u l t , o p e r a t i o n t h e

The problems enc ountere d dur in g th e BP-13 spa cec raf t checkout, and

th e anomaly i n the o pera tion of th e pwr~pduring th e countdown and f l i g h tl e d t o se ve ra l changes i n th e coolant pump fo r th e BP-:L~ pacecraf t .

The s t a in l e ss - s t ee l ba l l bear ings were i nd iv idua l ly in spec ted andselected for conformance to t he ABEC type 7 bear ing code t o main ta inthe i m p e l l e r shaf t a l inement .

r e s i n and baked under vacuum t o preve nt ox ida t ion or cor rosion of thes t a t o r . I n a d d i t i o n , t he n o t o r a nd pumps were assembled i n a dry n i t rogen

atmosphere, hermet ica l ly sea led , and sh ipped and s tored i n a dry n i t rogen -

filled sea led con ta ine r t o p r even t ox ida t ion ( f i g . 4 .12-3) .

The motor stator w a s sprayed with epoxy

U N C L A S S I F I E D '




Performance.- The equipment co ol in g subsystem was operated as .

planned during the launch countdown.

a t a ra te of2.8

ga l lons per minute t o the coola n t s to rage tank . Theonboard coolant pump and cabin fa n operated w i t h no apparent malfunction.

The coolant pump-outlet pressure was 31 p i g , indicating normal pumpoperation. The low pump-outlet pre ss ure and t h e h ig h t o t a l c u r re n t

measured dur ing th e BP-13 spacecraft countdown were not ev ident f o r the

BP-15 sp ac ec ra ft . During th e countdown f o r t h e BP-15 spacecraft, manual

programing of the operation of th e f a n and pump w a s used t o maintain tk-e

cab in -a i r and co ldp la te ou t l e t temperatures w i t h i n t h e range des i r ed

( f i g . 4.12-2).take n immedia tely pr io r t o umbi l ica l d isconnect are shown i n t ab le 4.12-11.

The GSE c i r c u l a t e d 16" F coolan t

The values of the equipment-cooling subsystem para met ers '

The telemetry data ind ica ted that the cab in f an w a s turned off by

the barawsi tches at T+334 seconds, as sh am by the d rop i n the e l e c t r i c a l

system t o t a l c u r r e n t f r m 42 t o 35 amperes. The ca bin pre ss ure a t t h i stime was 5 . 3 ps ia , which i s withi n the spe cif ied range and i n agreement,

w i t h prelaunch tes t ing .

The cabin pressure ra te of decay incr ease d a t n-161 seconds a t the)

tim e when th e launc h-esca pe subsystem ( U S ) xp los ive bo l t s were f i r e d

( f i g . 4 . 1 2 - 4 ) .

The BP-15 spacecraf t d id not inc lude gas-pressure- re l i e f o r con-

t r o l valves, and during launch, pressure was r e l i eved by leakage through

th e press ure sh el l . Cabin leakage was t o be equ iva len t to the flow

from a 4- t o - -inch d iameter ho le .

de ta i l r ega rd ing t h i s change i n cab in p ressu re decay rate . There wasno s t ruc tu ra l con f igu ra t ion s imi l a r i ty be tween the bo i l e rp l a t e cab inand the product ion spacec raf t cab in pressure sh el l .

1 1

2See sec t ion 4 . 7 for a d d i t i o n a l

The t e l eme t ry da ta i n d i c a t e d t h a t t h e c o o l a n t pump continued t o

opera te dur ing the launch and o rb i t a l phases or' the miss ion un t i l powerdep le t ion w i t h no apparent malfunction. The el e ct r i c al equipment

chass i s temperatures ranged from 43' t o 50" F a t launch and from 50" t o55" F a t the end of t h e f i r s t o r b i t a l pass.

The equipment cooling subsystem sa t i s fa c t or i l y fu l f i l le d thefunc t ion r equ i r ed of t he subsystem during the mission.




U N CLASS IF1ED4-214

TABLE 4.12-1. MONITORED EQUPMENT COOLING SUBSYSTEM PARAMrmERS

~~

Ground support equipment

Cabin air temperature

Coldplate inlet coolant temperature

Coldplate outlet coolant temperaturc

Tank outlet coolant temperature

Tank inlet pressure

Pump outlet pressure

GSE (S14-052 delivery t,emperature

GSE (S14-052) delivery rate

Telemetry

Cabin air temperature

Cabin interior pressure

TM RF transmitter A temperature

TM RF transmitter B temperature

TM RF transmitter C temperature

TM RF amplifier A temperature

TM RF amplifier B temperature

TM RF amplifier C temperature





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5.0 SA-7 LAUNCH-VEHICLE DESCRIPTION AND PERFORMANCE

7.1 Description

The Saturn I is a two-stage launch vehicle consisting of stagesS-I and S-IV, an instrument unit (IU), and various fairings and adapters.The total vehicle length is approximately 190 feet, consisting of an

80.3-foot-long by 257-inch-diameter S-I stage, a 41-foot-long by 220-inch-diameter S-IV stage, a 4.8-foot-long by 154-inch-diameter instrument unit,

and a 64.1-foot-long by 154-inch-maximum-diameter oilerplate spacecraftand launch escape subsystem (LES).presented in figure 7.1-1.

.

Vehicle details and dimensions are

The S-I stage dry weight is 107,170 pounds with a propellant capac-

ity of 850,000 pounds (lox and RP-1). Eight H-1 engines mounted in twoclusters, four inboard and four outboard, produce a total sea-levelthrust of 1.5 million pounds.

The S-IV stage dry weight is 13,857 pounds with a propellant ca-pacity of 100,335 pounds (LHthe S-IV stage produce a comgined thrust of' 90,000 pounds.

and loxj. The six RLlOA-3 engines of

The instrument unit contains most of the flight control equipment,including the vehicle inertial guidance and control system and the air-borne hardware of six tracking and four telemetry systems. The IU alsohas an integral power supply and distribution system, cooling systems,and a gaseous nitrogen supply system. The IU begins to function prior

to lift-off to command S-I start sequencing and to maintain programhg,sequencing, and flight control t h r o u g h S-I and S - I V stage operation.

Vehicle telemetry systems are provided for each stage and the IU.These systems include.six airborne systems and one digital data acquisi-

tion system for preflight checkout in the S-I stage, three systems inthe S-IV stage, and four systems in the IU.

5.2 Preliminary Flight Performance

After 11 seconds of vertical flight, the launch vehicle began tor o l l to the proper flight azimuth of 105" east of north and completedthe maneuver at W25.7 seconds. At ~ 1 2 . 6econds, the preprogramedpitch-attitude profile was initiated and continued until T+136.3 sec-onds, at which time an essentially constant pitch attitude was main-

tained until the initiation of active guidance at 17.3 seconds afterseparation of the S-I and S-IV stages.

occurred at W147.4 seconds which was only 0.7 second later than nominal.The shutdown of the S-I stage



5 -2

The actual trajectory parameters as compared with nominal were at that

time about 122.6 ft/sec (37.2 r.i/sec) high in space-fixed velocity,

1.19 nautical miles (2.2 lun) high in altitude, and 0.55 nautical mile(1.02 km) greater in range.

Separation of the S-I and S-IV stages occurred at W148.2 seconds,

followed by ignition of the S-IV stage 1 .7 seconds later.

rockets functioned as expected and were successfully jettisoned. Fol-

lowing the initiation of closed-loop guidance, the vehicle was steered

into a nearly nominal orbit after S-IV shutdown, which occurred about2 seconds later than had been predicted.

A l l ullage

The overall performance of both the S-I and S-IV propulsion systemswas highly satisfactory, with all engine parameters being near predictedvalues. A l l four of the 3-1 retrorockets performed as expected. Thetotal S-IV burn time was 1.3 seconds longer than predicted, as comparedwith the expected 20 dispersion of fll seconds. The S-IV common bulk-

head pressure was steady at approximately 0.8 psia with no indication ofleakage.

Stability of the total vehicle during S-I stage flight was main-tained through utilization of the ST-124 stabilized platform andassociated hardware.

trol rate gyros in closed loop during this phase of flight. In addi-tion, the ASC-15 guidance computer was used for the first time to gen-erate the r o l l and pitch program for s-I stage flight. A preliminaryanalysis of flight parameters indicated that all guidance and control

hardware functioned in a satisfactory manner. Lateral load torquescaused by wind were very small, and, consequently, the stabilizing kor-

ques from the gimballed engines were small (0.50" deflection or less).The m a x i m u m wind component, which occurred in the yaw plane, was only

49.2 ft/sec (15 m/sec). Since the pitch program was designed for zero

winds, the angles of attack were very small (1" or less during maximumdynamic pressure). A small aerodynamic r o l l moment (approximately 3" roll

attitude) was experienced during S-I stage flight which was similar to

what had occurred on earlLer flights. Pitch and yaw disturbances during

separation were very smal:, but at the start of separation a disturbing

torque caused the roll rate to build up to a m a x i m of 2.3 deg/sec with

a corresponding r o l l attitude of /5 " about 3 seconds after separation.

Upon S-IV ignition, the corrective action of the engine actuators caused

a m a x i m rate overshoot to 6 deg/sec with very adequate damping.r o l l rate observed would correspond to a 1.1" misalinement of one ullage

rocket.

This was the first test of the ST-124 and con-

The

At guidance activation, the S-IV vehicle attitude was changed from

67" to 75" nose down withLn 21 seconds in response to adaptive steeringcommands. The lateral deviations accumulated during S-I burn required



5-3

maximum transients of 1.5'' and 6" in the X and Y commanded angles,

respectively. Such transients are not considered severe.

A preliminary analysis of orbital insertion conditions ( S - I V cut-

off plus 10 seconds) has indicated that the difference between theguidance computer value of space-fixed velocity and the value obtainedfrom orbital tracking was 6.2 ft/sec (1.9 /sec). The lateral velocitydifference obtained was 14.8 ft/sec (4.5 m/sec) . Such comparisons are

preliminary, and the uncertainty in these values is from 3 to 6 ft/sec.

Very low bending moments (about one-fourth of the magnitude experi-enced during mission A-101) were observed during the flight.magnitude can be attributed to the low angles of attack experienced. Nosignificant disturbances were noted during launch-escape subsystem (LES)jettison, but some first mode bending was observed after jettisoning.

The low

Base thermal and pressure environments for the S-I stage weresimilar to those experienced during mission A-101.

of the S - I V and S-I stages indicated slightly higher values than had

occurred during mission A-101.to the fas ter trajectory flown on mission A-102.

Skin temperatures

These higher values were probably due

Overall performance of the launch-vehicle telemetry instrumentationsystem was good with only 8 of 1,233 measurements having failed completely.

A complete detailed evaluation of the performance of the launch

vehicle is given in reference 7.



5-4 U W L A S S F I E P . -

Launch escape~ subsystem

Spacecraft

I

istrumen t unit1 .-., - 8

tI

S-IV stageI Ullage

i I ?

rockets

2 2 0 "

, Retro

rockets

190'

Figure 5.1-1.- Apollo mission A 4 0 2 space veh ic leshowing cu taway views of launch vehicle.

c c -U NC LA-SS F IE L



6-1

6.0 8CIJCWDING REMARKS

All of the spacecraft test objectives for the Apollo mission A-102

were fulfilled.

The compatability of the spacecraft with the launch vehicle wasfurther confirmed under launch and exit condition.

Satisfactory engineering data, covering designated parameters of

spacecraft environment for a Saturn V type launch trajectory, were ob-tained for use in verifying launch and exit design criteria.

The launch-escape motor together with the pitch-control motor pro-pelled the launch-escape subsystem safely out of the path of the space-

craft in a demonstration of the alternate mode of launch-escape towerjettison.

Flight data from the instrumented simulated R C S quad assembly dif-

fered from the values assumed for design criteria for the RCS. Addi-tional investigation and analysis will be necessary to complete the

flight data and design criteria.

Flight .data on spacecraft launch environment near the R CS quad A

was insufficient for verifying criteria. Additional investigation andanalysis will be required.

The flight trajectory of the mission provided the launch environ-

ment planned f o r the mission.

A l l spacecraft subsystems performed the functions required for asat actory miss ion.



9. A N C L A S S F I ED.

7 . 0 APPENDIX A

7.1 Prelaunch Operations

7-1

I n i t i a l checkout of t he BP-15 spacec raf t was accomplished i n th e

Apollo t e s t and ope ra ti ons (ATO) area a t t h e c o n t ra c t o r’ s f a c i l i t y a tDowney, Ca lif or ni a. Fi na l checkout termi nated a t Cape Kennedy, F lo r ida ,

wi th the launch opera t ion.

The major t e s t s and oper ation s performed on th e sp ace cra ft , o r i n

conjunc t ion wi th sp acecraf t opera t ions, were conducted i n accordancewi th t he d e t e i l e d Operat ional Test Procedures (OTP) . (See appendix B,s e c t i o n 8.2.) These procedures define the step-by-step t e s t operat ionst o be performed and t he response o r values requir ed f o r approval, where

app l i cab l e .Downey, California, (ATO) and a t Hangar AF and Launch Complex 37E3 fa -c i l i t i e s a t Cape Kennedy, Fl o ri da . (See se ct io n 8 . 2 . )

The OTP’s were used throughout the checkout operations a t

On March 6, 1964, th e command module, s er v i ce module, i n s e r t ,

ada pte r, and launch-escape tower were t r a n s fe r r e d t o AT 0 from the

manufac tur ing f a c i l i t i e s . The schedule of‘ miles tone events for the

BP-15 spacec ra f t dur ing t he AT0 period i s given i n f i g u r e 7.1-1.

The schedule of ground support equipment (GSE) modif ica t ions andv a l i d a t i o n a t Downey, California, i s i n d i c at e d i n f igure 7.1-1. BP-13

spacec ra f t GSE electronics equipment w a s used a t Cape Kennedy ra t h e r

than sh ipp ing add i t i o na l GSE electronics from Downey.

Individual subsystems t e s t s were performed from April 29, 1964,t o May 11, 1964. Subsystems w e r e programed for individual checkout

dur ing th i s per iod i n coordina t ion wi th modi f ica t ion work on the ve-h i c l e and assoc i a t ed GSE. BP-15 s p a c e c r a f t subsystems t e s t s d i f f e r e d

from those of BP-13 spa cecr aft i n th at no spe ci al measuring dev ices

(SMD) o r sp ec ia l adapt ive devices (SAD) were used dur ing thi s phase of

t h e BP-15 spacec ra f t ope ra t ions .

Stac king , alin emen t, and hookup too k pla ce between May 11, 1964,The spacecraft assemblies were mated and a l i n e d i nnd May 18, 1964.

the Navajo tower and a l l GSE w a s connect ed i n p repa ra t i on f o r t he i n t e -

gra ted subsystems t e s t . P r i o r t o t h e s t a r t of t h i s phase of opera t ions ,

t h e mass ch ara c te r i s t i c s were de te rmined f o r each space craf t module. A

comple te qu a l i ty cont ro l insp ec t ion by th e spacecraf t ’ con t rac tor ‘s and

NASA pe rs on ne l was a ls o accomplished.

The int eg ra te d subsystems t e s t w a s s t a r t e d on May 19, 1964, and

w a s completed with th e excep t ion of th e l a s t 10 s teps (e lec t romagnet ic



7-2

i n t e r fe rence (EM) e s ts ) of the procedure. The spa cec raf t equipment

w a s opera t ing normally, but th e pyro -subs t i tu te u n i t had grounding

problems s imilar t o those encountered during BP-13 spa cecr af t checkouta t Cape Kennedy. The in te g r a te d subsystems t e s t w a s continued on

May 25, 1964, ut i l i z i n g f'used, 1-ohm re s i s t o r c i r cu i t s t o measure f i r ingand/or t r a n s i e n t s i g n a l s . CEC rec ord ers were used t o read th e f i r i n g

indica t ions . The id en t i ca l t echnique w a s u t i l i z e d p r ev io u sl y a t Cape

Kennedy with the BP-13 s p a c e c r a f t . The EM1 por t i on o f t he i n t egra t ed

subsystem t e s t was then completed.

Spacecraft removal from the Navajo tower was performed followingdeb rief ing and acceptance of th e in te gr ate d subsystems t e s t . The ser-vice module, in se r t , adapter , and ass or t ed GSE were shipped on June 3,1964, by th e B-377 PG a i r c r a f t .

tower were s hi pp ed on June 19, 1964.

fo r t h i s shipment, because of the non-ava i l ab i l i t y of an Air Force C-133B.The spacec ra f t and a s soc i a t ed GSE were t r a n s p o r t e d t o Cape Kennedy with-

out damage t o t h e equipment.

The command module and launch-escape

The B-377 PG w a s also u t i l i z e d

.The BP-13 spacec ra f t ope ra t i ons a t Cape Kennedy, F lo r id a , began

wi th t he rece iv ing i nspec t i on of t he GSE and spacec r a f t a s sembl ie s i nHangar AF. P r i o r t o a r r i v a l o f t h e s p a c e cr a f t, t h e d e c is i on was made

that no hangar t e s t ing would be performed. This a l l e v i a t e d t h e n eed

for removal and r e t u r n of th e BP-13 s pa ce cr af t GSE from Complex 37B.Figure 7.1-2 re sen t s t he scheduled mi l es tones f o r t h e BP-15 spacec ra f t

pre launch opera t ions .

The command and service modules were mated i n Hangar AF on June 22,

1964. F i na l i n s t a l l a t i o n of t h e a d a p t e r ai r - c o n d it i o n in g b a r r i e r wascompleted, and t h e command and s er v ic e module assemb ly was mated t o t headapter. (See f igure 7.1-3. The ent i re s tack w a s then s ing l e -po in t

weighed and loaded on the ver t ica l t ranspor t .

t r anspor t ed (in he s t acked conf igura t i on) t o Complex 37B f o r mating

wi th the launch vehic le (SA-7) on June 26, 1964. Figure 7.1-4 shows

the spacec ra f t be ing ho i s t ed .

"he spacec ra f t was then

The instrumented RCS quad assembly was received from Downey,Cal i fornia , and i t was i n s t a l l e d and checked out immediately following

the mat ing of th e sp acecraf t wi th th e launch vehic le .

comple-ted on June 29, 1964.This task was

BP-15 spacecraf t pre launch opera t ions were extended approximately2 weeks because of a st re ss- cor ros ion problem i n the l i qu id oxygen dome

of the S-I engi nes. The ei gh t engines were removed, reworked by th ec o n t r a c t o r , s t a t i c f i r e d , a n d r e i n s t a l l e d i n the S-I stage. During this

time, a s p e c i a l t e s t w a s performed a t th e req uest of MIT, for th e



7-4 * -UNCLASSIF IED ,

The overall BP-15 space craf t pre launch opera t ion, both a t Downey

and Cape Kennedy:progressed smooth ly.

opera t ion tha t exper ience ga ined dur ing the pre launch opera t ions oft h e BP-l3 spacec ra f t was being ut i l ized and the methods were being

improved.

I t w a s evident throughout the

U CLASS I F ED e1



U NCLASS IF IED 7-5

I

I '




7-6 U N CLASS IF1ED

U N CLASS IF I ED



U N C L A S S IFSED 7-7

UNCLASSIFIED



Figure 7.1-4*- fsP- A-7 launch vehicle at



UNCLASSIF IED 7-9

Figure 7.1-5.- L ES being lifted for mating to the spacecraft& Launch Complex 376,Cape Kennedy, Florida.

U N C L A S S IFIED




7.2 Launch Operations

The T-1 day countdown began a t7:30

a . m . e . s . t . on September17,1964, a t T-875 minute s, The sp ac ec ra ft po rt io n of th e countdown, how-

e v e r, d i d n o t b e gi n u n t i l 9:40 a . m . e .s . t . and consisted of tower-bol t

o rd na nc e e l e c t r i c a l c on ne ct io n cl o se o ut a n d e l e c t r i c a l v e r i f i c a t i o n .

The count w a s resumed a t T-12 minutes and progress ed t o T-5 minuteswhen a Range s a fe ty hold w a s ca l l ed because o f in t e rmi t t en t ope ra t ion o f

t h e Grand Turk I s l a n d radar. Because of S-IV lox bubbl ing and spacecraf t

b a t t e r y l i f e t ime c o n s t r a i n t s , the count w a s r e c yc l e d t o T-13 minutes, and

t h e s p a c e c r a f t was t r ans f e r r e d t o ex t e rna l power.

c u l t y was encountered wi th t he swing arm hydrau l i c t e s t . This problem

During th e ho ld , d i f f i -

The usual T-1 day activities performed on t h e BP-15 spacecraft were

conducted on T - 1 day of the countdown demonstration t e s t on September 1 4 ,1964, h u s e l i m i n a t i n g t h e n eed t o r e p e a t t h e s e t a s k s .

The f i n a l countdown began a t ll:25 a . m . e . s . t . on September 17,

1964, a t T-545 minutes , wi th a planned T-0 ime of 1O:OO a . m . e . s . t .Figure 7.2-1 shows the schedule o f ac tu al and p lanned t i m e f o r t h e t a s ksperformed during t he launch count. T he spacec ra f t t e s t ing p roceeded

normally without any holds o r equipment malfunctions. Only two eventsof major s ign i f i can ce occurred dur ing the count :

encountered by the Range and Hangar S ground s t a t i o n i n dis t in gui sh ing

t h e C-band beacons responses because of multipath interference from theP a t r i c k A i r Force Base r adar , an d t h e o t h e r w a s an inadver t en t ac tua t ion

a t T-360 minu tes o f th e se rv ice s t ru c tu re ad j u s tab le &- leve l F i r ex sys temduring removal of a n a i r con diti onin g duct . Approximately 50 percen t o f

the serv ice module e x t er io r was wetted.

one was t h e d i f f i c u l t y

The count w a s continu ed u n t i l T-245 minutes when a hold w a s c a l l e d

because of the w e t umbilicals. The water had entered one S - N u m b i l i c a l

connector which, i n tu rn , produced erroneous ind ica tio ns of S-IV engine

e x c i t e r firing. Power was removed from t h e S-IV s tag e , and th e mois ture

w a s d-r ied from t h e conne ctor. The count was resumed 69 minutes l a te r .

The count cont inued unt i l T-3O minutes when a scheduled hold began.

Analysis of t h e problem

During t h i s scheduled 21-minute hold, t h e S-IV l iquid oxygen ( l ox ) pre -

p r e s s u r i z i n g r e g u l a t o r i n d i c a t e d a malfunction.

i n d ic a t e d t h a t th e r e g u l a t o r w a s opera t ing sa t i s f ac to r i l y ; however, t h e

hold had been extended 4 minutes longer th an scheduled. The count pro-

g r e s s e d t o T-12 min ute s when i t w a s again in ter rup ted because of a m a l -

funct ion ing S-I hydraulic-pump temper ature in te rl oc k, which preven ted t h e

S - I hy dr au li c pumps from being s t a r t ed . Sin ce measurements ind ica ted

normal temperatures, th e in t er l ock w a s j m p e r e d ou t i n t h e blockhouse

d i s t r i b u t o r . The t o t a l h ol d t i m e w a s 20 minutes.





w a s cor rec ted by a jumper i n a blockhouse d is t r ib u t or wi thout adding t cthe Range hold. A f t e r 50 minutes, th e ra da r problem w a s corrected, and

the count resumed and was continuous through l if t-off which occurred a t

11:22:43 a.m. e . s . t . on September 18, 1964.





-'F si1e.wIIsi lence

Nr 011

-

r

preps I;e crew report in

d crew to statio i l (adj t istable 5)

I ICoiii icct twr leqs and close out

I

I I

1LE S i gn i te r *

/2 c lose ou t

I l l s tLafety

-bal l p i

f

-

-S eiiab

-ey to 11

D -I

I I

r ies

1- ctualI0 l a y e d

Performed 011 Sept 1,

inned hold

nd no t reoeated ,

870 840 810 780 750 720 690 660 630 600 570 545T-time, iiiin

I I

07:30 p.m. Q8:30

I I 1 I I09:30 10:30 11:30 .m. 1230 a.m. 1:30

' ( S e p t 18)Sept 17 e.s.t., hr:min

(a) T -1 day, September 17, 1964.

Figure 7.2-1.- Apo l lo miss ion A-102 countdown act iv i t ies .

U N CLASS 1 F I ED




5 2-508

rL

I I

isc xdcr & remove hatch

- ctual

. En d 10 hr 30 min hr >T -545

40 510 480 450 420 3

T-time, min

Fina l insp IInstall hatch i

Vacturn

Ladder ou t

11:30 p.m. 12:30 a.m. 01:30 02~30 03:30 04:30Sept 17 Sept 18 e.s.t. , hemin

(b ) Launch day, September 18, 1964.

Figure 7.2 -1 . - Continued.




7-14 U N C L A S S IF l ED

r

I I I H& for i

I I

IServ ice structure moving

I I I I

1n S C p w r o n RI

C-band

I I I I I I

3 210 180 150 120 90 60 30 30

T-time, min

-i

?

--12 I

-&Hold for ii 20 min

I

*i Hold for 50 ii min, recycle i

i t o T - 1 3 i

5 13 0

:-band on

TM RF on

Nr txfr

l i f t -of f

I I 1 I I I I

07:30 08:30 09:30 10:30 11:3005:30 a.m. 06:30

Sept 18e.s.t. , hr:min

(c) Launch day, September 18, 1964.

Figure 7.2-1.- Concluded.

U N CLASS IF1ED




which s ign a l t o t r ack , bu t s i nce t he second re tu rn ( spacec raf t t r an s -

ponders) was t h e s t r o n g e r s i g n a l , i t was t racked ins tead of the IUt ransponder. See se ct i on 9.1.

( 2 ) Fgl in A i r Force Base, F lo r id a , on t h e f i r s t o r b i t a l p as s, was

reques t ed t o s k in t r ack r a the r t han i n t e r r oga t e t he beacon which w a s

s t i l l ac t i ve . Sk in t r ac k w a s not obtained because of t h e extreme range.

( 3 ) White Sands Missile Range, New Mexico, te leme try sta t i on , on

t h e f i r s t or b i t a l pa s s , r epor t ed heavy mul t i pa th i n t e r fe rence on l i nk A

( t h e acqu i s i t i on l i nk ) between 0" and 25" e l eva t i on .

vari ed al so , poss ibly from spac ecraf t tumbling.cons tant signal s t reng th dur ing t he second orb i t a l pa s s .

The s i g n a l s t r e n g t h

Link A had a f a i r l y

(4) Pre tor i a t e l eme t ry s t a t i on repor t ed dropout s of 4 seconds and

13 seconds durat ion on t h e t h i r d o r b i t a l p as s an d l a t e a cq u i s i ti o n a n da 15-second dropout of l i n k C on the fourth o r b i t a l pass . Links A an d

B were s a t i s f a c t o r y .

(7) P r e t o r i a t el e m et r y s t a t i o n , o n t h e f i r s t o r b i t a l pas s , w a s

unable t o hold automatic t r ac k te lem etry due t o undetermined equipment

problems. I n swi tching t o manual t rac k, no dropouts i n recept ion w e r e

experien ced; however, high n oi se was observed i n some per iod s of t h e

t rack.

(6) Grand Turk I s l and radar s t a t i o n l o s t c o n t a c t w i t h t h e v e h i c l e

from 00:06:43 t o 00:07:55 g.e . t . because of Bermuda interference onbeacon reception.




U NCLASS I F IED 7-17

c

"-




7-18

h q u i s i t i o l i lass ncqui si i o n LOSS

U N CLASS IF I ED

A cqui s i t i on Loss

~

Cape Kennedy 03: 00: 00

Patrick A i r Force Ease, 0 0 :0 0 :1 6Fla.

w a n d B a r n I s l and 00 :01:14

ssn Salvador I s l a n d w:02: 5

Gr and T u r k I s l a n d 03: 03

Antigua Island

Ascension Island 0021:10

Pretoria, South Africs 00 : 32: 25

;

BeImUd8 00: b:28 00: 11:k 3

Carna-on, A ust ra l i a 00: 54 : 19 00: 9: b6

Depanment of Defense Range

Hawaii 01:lR29 01:21:29 02:53:57 02 :56 :29 04 :25 23

C a l i f o r n i a 01:26:33 01: 53:27

White Sands Missile Range. 01:31:35 01: 6:23 05: 5: 23 03: 08: 5

N. Mex.

00: 9: 16

00:07:h5

00 :09: 58

00: 10: 53

am:11: 5

S t a t i o n s

a:31:11

01: 6: I:,

01:j8:38

O1:4:: 58

01: 55: 15

02:&:01

Manned Smce P l iP h t Wetwork




8

U N CLASS I F ED

7 .4 Data Coverage and Avai l ab i l i ty

7-19

Data suppor t , genera l . - All da ta reques ted by the Apol lo Spacecra f t

Trogram Office, Test Evaluation Branch, MSC-Houston, f o r t he e va lua t i onof l aunch vehic le and spacecra f t pe rfomance are l i s t e d i n t a b l e 7.4-1.

Re spons ib i l i t y f o r a c qu i s i t i on and de l i ve ry of t h e se da t a was

di vi de d between th e Kennedy Space Cent er (KSC) and the Goddard Space

Flight Center, Florida Operations (GSFC-FO) .d a ta and GSFC-FO provi ded in se rt io n- ph as e (Antigua ) and orbi ta l-phase

data . Both Centers del ivered a l l data t o t h e Data Sect ion, Operat ions

Sup por t Branch, MSC-Florida Op era tio ns (OSB) f o r h an d li ng a nd d i s t r i -

bu tion.

KSC pr ov ided laun ch-phase

Telemetry osci l lograph recordings were annotated a t OSB p r i o r t o

d i s t r i b u t i o n t o t h e l au nch s i t e a na lys i s and re por t i ng team. Real-timet e lemet ry osc i l lograph recordings were made av a i la b le t o the team wi th in

hours a f t e r l i f t - o f f .

h u n c h d a t a s up po rt . - Delivery of quick-look da ta requi red by th e

a na lys i s a nd re por t i ng team w a s sa t i s fac tory . Magne t ic t apes from

!Telemetry Building 2 (Te l 2) , Telemetry Building 3 ( T e l 3 ) , Mission

Control Center (MCC) , and Hangar S were all rece ived wi th in a t ime

envelope of 1 o 4 hours ( tab le 7.4-1).from e1 2 were rece ived wi th in 1 hour a f t e r l i f t -off . Quick- look

processed da ta p lo t s (SC-4020 plots) of commutated and continuous

channels were received 24 hours a f t e r l i f t - o f f . Telemetry s ignalst re ng th re cord s were receiv ed between 5 and 6 calendar days a f t e rl i f t - o f f . D e l i v e r y of engineer ing sequent ia l f i l m i s l a t e a t t h i s .

wri t ing and thus prec ludes a complete eva lu at io n of photographic coveragea t t h i s t i m e .

Real-time osc i l lograph records

I n s e r t i o n and o r b i t a l da t a support . - A 24-hour delay w a s experiencedi n obtai ning ins er t io n phase te lemet ry tap es from Antigua. The delay

w a s a t t r i b u t e d to breakdown of t h e Air F orce a i r c ra f t s cheduled t o t rans-port data from Antigua t o P a t r i c k Air Force Base.o r b i t a l t a p es provided by GSFC-FO were ge ne r a l l y p rov ided wi th in t he

time f rames spec i f ied .

The msjo r i t y of

Orbi t a l phase tapes from H a w a i i , Bermuda, Corpus Christi, and

Eg l in were received at GSFC, Gre en be lt , Maryland. Copies were forwarded

d i r e c t l y t o MSC-Houston and ar ri ve d a t Ti-14 alendar days. A t t h e

present t ime, t h e tape from Pretoria has not ar r ived a t GSFC.l a y was caused by breakdown of successive A i r F orc e a i r c ra f t e n rou t e

t’rom Pr eto ri a t o Washington, D. C.

The de-




7-20 UNCLASS IF I ED

Photographic coverage. - Photographic coverage, including th equan t i ty of instrumentation committed and data obtained during the

launch-phase, i s shown i n ta bl e 7.4-1. The t o t a l r eq ui re me nt s f o rengineering sequential camera coverage and usage are Shawn i nt a b l e 7.4-11. Locations of the engineering sequential cameras areshown i n f i gu re s 7.4-1 nd 7.4-2.

An evaluat ion of a l l films r eceived ind ica te s th a t , of the s i xtracking camera f i l m s ava i l a b le fo r review, t h e q u a l i t y was g e n e r a l l ygood wi th resp ect t o exposure, focus , and t rack ing , except fo r one f i l m .

Fi lm qu al i ty wi th r espec t t o o the r camera and f i lm def ec ts was good wi th

the fo l lowing except ions:

provided no usable timing image.

foc al- len gth tr ac ki ng cameras and provided ve ry good coverage fran

approximately 20 seconds a f t e r l i f t - o f f t o approximately 32 seconds

aFter LES j e t t i s o n (~160.2e e ) .en t i r e space veh ic le t o s t ag ing and then followed t h e S-I s tage .

Cocoa Beach and Melbourne Beach ROT1 cameras t racked the en t i re spaceveh ic le through S-I staging and t h e n t h e S - I V s tage and the BP-15spacecraf t th rough LES j e t t i s o n . The Vero Beach ROT1 tracking cameral o s t t r a ck of t he spacec ra f t a t LES j e t t i s o n b u t d id provide goodcoverage of t he spacec ra f t p r i o r t o and fo llowing tower sepa ra t ion .

The Cocoa Beach, Melbourne Beach, and Vero Beach fi lm s d i d pro vid e

s u f f i c i e n t data t o de termine tumbl ing rates and approximate tr aj ec to ryo f the j e t t i soned M S . The Grand Bhama Island IGOR camera providedno usable data on th i s miss ion .

One f i lm appeared ver y grainy, and two f i l m s

Five of the films were from the long

The Patrick IGOR camera tracked the

The o the r f i lm available for review provided a close-up view or'

t h e s p a c e c r a f t a t l i f t - o f f and fo r app rox ima te ly 35 seconds of e a r l yf l i g h t .

MSC-Houston dat a redu ctio n.- Spa cecr aft tele me tr y da ta were pro-

cessed a t MSC-Houston by the Computation and Analaysis Divis ion (CAD),

The tape copy used f o r launch-phase d ata re duc tio n (T-10 sec t o approx-

imate ly Who0 s e n ) was from the Mission Control Center (MCC) telemetry, s t a t i o n , Cape Kennedy, Flsi., and w a s r ece ived in Houston a t W23 hours.

The tape copy used for reduct ion of inser t ion-phase da ta (W400 se e

th rough in se r t jon ) w a s from Ant igua I s land te lem etry s ta t i on and w a s re -ceived i n Houston a t a p p r o x i i a t e l y "72 hours. The f i r s t package of

eng inee r ing ana ly s i s p lo t s f rom these da ta t apes was p ro vi de d t o t h eevaluation team a t ~ i - 6 alendar days.

h i s t o r y p l o t s and t s b u l a t i o n s of a c c e l er a t i o n s , s t r a i n g ag es , e l e c t r i c a l

root-mean-square (RMS) lo t s of low-frequency acceler a t ion s , f l uc tua t i ngp ressu res , and r a d i a l v ib ra t ions . Con ical su r face p ressu re coe f f i c i en t s

were determined by using th e measured c on ic al su rfa ce pre ssu res and the

with support f ram Instrumentation and Electr ical Systems Division (IESD).

me package contained time-

funct ions , tempera tures , hea t f lux, con ica l su r f ace p ressu re coe f f i c i e n t s ,





dynamic pressure based on the measured atmospheric den si t y a t the t ime

of launch.

a v a i l a b l e w i t h i n"9

calendar days and contained one-third octave-band

time hi s t or ie s ; power spe ct ra l de ns i t i es (PSD) ; and one-third octave-

band sp ec t r a l ana ly ses of f luc tua t ing pressures , r a d i a l v ib ra t ions ,

s t r a in gages , and acce le r a t ions .

A second package of engineering analysis p l o t s was made

The RMS, PSD, and one -thi rd octave-band ana lys i s pl ot s were pro-

duced using the MCC tape. Reduction of vi br at ion da ta was accomplished

by both analog and d i g i t a l methods. The PSD p l o t s w e r e produced on t h e

d i g i t a l computer and both RMS, and one-third octave-band an al ys is pl ot s

were produced us ing ana lo g equipment.

L i f t -o f f ( T - 0 ) w a s es tab l i shed as 11:22:43.26 . m . e . s . t . as es tab-

l i shed by launch vehic le m b i l i c a l disconnect (17: 22: 43.26 G. m . t . ).

Event times were presented on the plots. The event times ind ica ted weret he b e s t information avai lab le a t the t ime of prepa ra t ion of t he p lo ts .

Prelaunch R and Z (Range and Zero) cal ibra t io n va lues , which vere re-corded for the continuous and high-level commutator parameters, were

wi th in 2 percent of t he or ig ina l va lues .r e s u l t or' R and Z ca li br at io n changes because th e change t o th e da ta

vould not have been signif icant .

the low-level commutator were not pre ci se measurements; c onsequen tly,they could not be used f o r d a ta reduction.

No co r r ec t ions were made as a

The R and Z cal ibra t ions recorded f o r

The number and type of parameters processed f o r BP-13 spacecraf tare compared with parameters processed f o r BP-13 spa ce cr af t i n

t a b l e 7.4-111.

All processed pressure measurments were b i a s e d t o read ambientDresswe a t launch. The ac ce le ra ti on s were bias ed to read lg on the

X-axis and zero-g on t h e Y- and Z-axes.

The data were processed and presented i n accordance w i t h require-ments es tabl i shed j o i n t l y by t h e Apollo Spacecraf t Program Office and

the a na ly s i s and r epo r t ing team subsystem anal yst s p r ior t o the f l i g h t .

Accelerat ion and strain-gage t ime-history da ta were processed and

p resen ted i n two d i f f e r en t fo rma ts .changes of 3 percen t , o r g r ea te r , of t el eme t ry f u l l sca le was superimposed

on a basic sample r a t e of one point every 0.1 second. A second method

u t i l i z e d t h e s a m e e d i t rou t ine a f t e r a 21-point running average1 had

t e e n performed.

An e d i t r ou ti ne t o determine

Data recorded by the hi gh- lev el commutators, except f o r RC S param-,eters, were processed u s ing an ed i t r ou t ine t o de te rmine t r ans ien t s




7-22 U NCLASS IF I ED

g r e a t e r t h a n 2 percent .

r m t i n e were superimposed on a basic sampling ra te of one every - second.

These data were plot ted and printed.

The data produced from t h e t r a n s l e n t d e t e c t i o n1

2

1The RCS temperatures were processed wit h a ?-point running average

and eve ry po in t p lo t t ed .

wi thout perf omi ng a running average.Every commutated data poin t w a s tabula ted

Every data poin t recorded by the low-level commutator was processed,

p lo t t ed , and p r in t ed .

calorimeter and body temperature data , but no e d i t rou t i ne w as performed.

A ?-point running average w a s performed on the

An appr oxim ate 3-second drop-out i n the da ta occurred on a l l re-

c e iv i ng s t a t i o n t a p e s j u s t p r i o r t o and during launch-vehicle s tag ing

because or' flame a t tenua t ion f rom the S-I sta ge re tr or oc ke ts. The samedrop-out occurred on t h e BP-13 spacecraPt f l i g h t .

MSC-Houston da ta processing anomalies.-

a. The MCC tape, used for pr imary da ta reduc t ion a t Houston, didno t ca r ry t he 5 - s t ep VCO ca l i b ra t i o n requ i red fo r da t a p rocessing a tHouston. A s a r e s u l t , it w a s necessary t o u t i l i z e t h e ground s t a t i o nbzmi-edge ca l ibra t ions recorded on the MCC t ape . The adequacy of t h i snrocedure w a s v e r i f i e d by comparing R and Z ca l ib ra t ion s recorded on the

Hangar S t ape wi th R and Z ca l ibra t ion informat ion recorded on the MCCt ape .

system accuracies, and no bias was present .

The values or' t he se R and 2 ca l i b ra t i on d a t a were well wi th in

b. The Hangar S tape was not used fo r , da t a r educ t i on a t Houston

because the t ime information recorded on t h e t a pe w a s di rec t - recordedas p a r t of a composite.

information recorded on the t ape incompat ible w i t h exis t ing ground-s t a t i o n equ ipnent s i nce t h i s equipment requires that time informationbe recorded separa te ly as ZL frequency modulated su bc arr ier .

This procedure rendered the ext rac t ion or' time

e. The tape from Tel 2 was not used because the tape containeda dropout of approxlmately 5 t o 8 seconds a t approximately Wl3O secondsvhiclh w a s not apparent i n any of t he o the r t e l eme t ry t apes .

A running average i s accomplished by averaging 21 consecutive ,

1

poin t s and p r in t i ng t h i s va lue ve r sus t h e t ime corresponding t o the

rr,iddie o r e leventh point . The number one data point i s dropped, anew o r 22nd included, and a new average i s produced and printed out

irersus t;fie new midpoint. This i s c o nt in u ed u n t i l a l l t h e data poin t shave been inzluded i n an average .

U NCLASS IF1ED



U N CLASS1F I ED 7-25

d. The 100-kc refe re nc e frequ ency on t h e tap e copy from Tel 3 used

a t Houston w a s approximately 1 .7 kc high.

Mzrshall Space Flight Center data. - Three s pa ce cr af t measurementswere t ran smi tte d through the launch-vehicle telemetry system on a sing:-eside-band link. D s t a from the three measurements were processed a tMarshall Space F l i g h t Ce nt er (MSFC) and forwarded t o MSC-Houston.

measurements are the folluwing:The

Launch-Iaunch- vehicle

telemetrypacecraf t vehi l emeasurement measurement chann el

Adapter radia l v i b r a t i o n no . 5 AAO@?D E3 3-900 S3 -13 E 01

Adapter ra d i a l v ibra t io n no. 6 AAOOgOD E374-900 S3-14Serv ice module ac ou st ic ~ 7 6 0 ~ 69-900 53-12

L u n ch - ve h i cl e d a t a e s s e n t i a l t o t h e a n a l y s i s of spacecraf t per -formance were processed a t MSFC and forwarded t o MSC-Houston.

vehicle parameters processed f o r MSC-Houston a re l i s t e d i n t a bl e s 7.4-1Vand 7.4-V.

L3unch-

All MSFC processed data were furn ished i n formats and engineering

un i t s compat ib le t o MSC analyses .




7-24

Presentationt a type


TABLE 7.4-1. - DATA AVAIIABILTI'Y

Requested delivery Date/time received

(a) (a1

I T+2 CD

T+2 CD

I Telemetry data from magnetic tapes

Telemetry Building 2


Mission Control Center

Hangar S

Hangar D

Grand P a h a Island

Antigua

Ascension

Eglin

Retoria

San Salvador

Hawaii

Bermuda

California

GuaYmaS

Postlaunch instrumentationnessage

L-in. magnetic tape2

5-in. magnetic tape

--in. magnetic tape

?-in. magnetic tape

2

h n . agnetic tape

&-in. magnetic tape

2-in. magnetic tape

Ll n. magnetic tape

A-in, magnetic tape

L-in. magnetic tape

i-in. magnetic tape2

1-in. magnetic tape

1-in. magnetic tape

1-In. magnetic tape

1-In. magnetic tape

Report (Iwx)

2

2

2

2

2

2

T+4 H

T+ 4 H

T+4 H

T+1 H

T+2 CD

T+ 2 CD

T + l 2 H

T+ 2 CD

T+2 CD

T+ 2 CD

T+ 2 CD

T+ 2 CD

T+2 CD

T+ 2 CD

T+2 CD

T+l CD

T+l H

T+!3 H

T+3 H

T+8 H

Did not acquire

T+1 CD

T+2 CD

T+4 CD

T+ 4 CD

(b)

T + l CD

T+14 CD

T+14 CD

T+14 CD

T+14 CD

T+l CD

Data on telemetry receiving station performance

I I I

Signal strength

Telemetry data sheets(station logs)

RECO r o l l

Sheet

T+(5 to 6)m

T+ 5 CD (partial)

Data from oscillograph

I I ITelemetry Building 2


Real-time osc llo-graph roll

Pleyback copy

T+l H

n S H

T+l H

~ + 1 8

aKey:

CD - Calendar Day WD - Working Day H - H ~ W S

bData requested but not received during the postlaunch reporting period.





Dsta type

Hangar S

Hangar D

7-25

Presentation Requested de liv ery Date/time receive da ) (a 1

Real-time os ci llo - T+2 H T+3 H

graph

oscillograph'Spe cial real-tim e T+IZ R T+3 CD

-

SC 4020 pl ot s -roll copy

SC 4020 pl ot s - book

T+ 5 H "1 CD

T+10 H T+l CD

TABU3 7.4-1.- DATA AVAILABILITY - Continued

Continuous channels

Continuous channels

!C+l D1 T+2 CD Ttl.5 CD

SC 4020 pl ot s - film

SC 4020 plots - book

~-

Commutated channels ISC 4020 plot s-fi lm I T+5 H 1 W l D

Conmutated channels

Conmutated channels

A l l launch vehiclemeasurements

A l l launch vehiclemeasurements

SA-7 measurement program

SA-7 ground and environ-menta l measurement program

Flash reports from groundand environmental measure-ment program

Antigua

Antigua

Ground and environmentalmeasurements

bunch-vehicle data

S C 4CQO p l o t s - book

Calibration curves -books

B x k

Book

Pase

SC 4020 plot6 - book

SC 4020 pl ot s - book

SC 4020 plots

T+2 CD

T-0

T-14 CD

T-14 CD

As available (notto exceed T+b CD )

W3 CD

T+b CD

T+b CD

T+i CD

T- 3 CD

T-14 CD

T-14 CD

(b1

T+3 CD

T+4 CD

T+3 CD

*Key:

CD - Calendar I ky WD - Working Day H - HOWS

bTsta requested but not received dwing th e postlaunch reporting period.

U NCLASS IF IED



7-26

Presentationata type


Requested delivery Date/time received

(a 1 ( a )

!CABLE 7.4-1.- DATA AVAILABILITY - Continued

Quick-look Ip transforma-t ion (pos i t ive veloci tyaccelerat ion)

Quick-look Ip t ransfoma-t ion (pos i t ive veloci tyaccelerat ion)

Quick-look IP transforma-

t ion (pos i t ive veloci tyaccelerat ion)

Quick-look Ip transfonna-t ion (pos i t ive veloci tyaccelerat ion )

Best estimate oft r a j ec t o ry

Aerodynamic parameters '

containing dynamic pres-sure, Mach number,Reynold's number/foot

Aerodynamic parameterscontaining dynamicpressure, Mach number,

Reynold's number/foot

F i nal a t t i t u de da t a(yaw, pi tch , roll)

Final reduced fl ight t e s tdata (coordinate system 1'

Final reduced fl ight testdata (coordinate system 1:

Final reduced fl ight t e s tdata (coordinate system 2:

Final reduced fl ight t e s tdata (coordinate systmo 2:

Final reduced flight t e s t

data (coordinate system 3:

aKey :

CD - Calendar Dsy

Magnetic tape

I I M printout

lhgnetic tape

IRM pr in tout

IRM pr in tout

I€&l r in tout

Magnetic tape

IEM pr in tout

Magnetic tape

E?&?r in tout

Magnetic tape

IBM printout

Magnetic tape

T + 6 H

T 6 H

T+3 CD

T + 3 CD

T +3 CD

T+ 3 CD

T +3 CD

T+ 3 CD

T+ 3 WD

T+ 3 CD

T +3 WD

T+3 CD

T+3 WD

WD - Working h y

T + l CD

T+1 CD

T+ 1 CD

T + 1 0

(b 1

bData requested but not received during t he postlaunch reporting period.

H - H O U ~ S





TABLE 7.4-1.- L W A AVAILABIUFY - Continued

mta type

Final reduced f light testdata (coordinate system 3 )

Flnal reduced fllght tes tdata (coordinate system 4 )

Final reduced f light tes tdata (coordinate system 4)

Position, velocity andspec ia l t ra jec tory para-

meters ( o r b i t a l )

Position, velocity andspec ia l t ra jec t ory para -meters (orbital)

Preliminary estimate ofdata coverage

Plotting board cherts( s ta t ions 1and LCC 37)

Plotting board cbarts(downrange stations)

Orbital f l l gh t parameters

Radar beacon log (uprangeETR s i t e s )

Radar beacon log

(danu9nge

(uprange 1

Radar event record (DOD'

s i t e s )

Fledar event record(NASA s i t e s )

Radar function record(includes signal strengthETR s i t e s )

Radar event record

Presentation

IBM pr intout

Megnetic tape

DM pr intout

D l pr in tout

Ikgnetic tape

Sheet

Sheet

Sheet

I T Y message

Sheet (log format)

Sheet (log f onm t )

Str ip char t

S t r ip chart

Str ip char t

S tr ip char t

Requested delivery

(a 1

93 3 CD

T+3 WD

T+3 CD

9 1 5 CD

"15 CI)

T+2 H

T + l H

T+2 CD

T+b H

T+ 1 CD

!r+4 CD

T + l CD

T+6 CD

T+6 CD

T+l CD

CD - Calendar b y WD - Working Da y

bmta requested but not received during th e postlaunch re portin g period.

B - H o ~ e

hte/time received

(a 1

T+6 CD





TABLE 7.4-1.- DATA AVAILABILITY - Continued

7-29

PFiD

Page

220.6

220.8

220.8

220,.8

220.9

220.13

220.13

220.13

220.13

220.10

220.10

220.10

'Key:

00I t eno

2

--

1

2

3

2

2

3

4

4

1

1

1

Data type

35 mm from T- 5 sec toT+30 sec

35 mm from T-5 sec to LQV

35 nun from T-5 sec to LOV

35 mm from t- 5 sec to LOV

35 mm reduction print froir-5 sec to T+ 4 min

16 m from T-30 sec tor+5 se c

16 mm from T-5 sec toP+lO se c

55 mm from T-5 sec to LQV

16 mm from T-5 sec to Lo V

'0m from acquisi tion

;o Lov

15 mm from acquisit ionso Lov

.6 m frcm acquisition80 LOV

Presentation

1 camera t o showvehicle motion forf i r s t 15 fee t off l i g h t

2 cameras i n tandemsharing ent ire vehi.cl e and launcher

2 cameras in tandemshoving entire vehi-cl e and launcher

2 cameras in tandemshoving entire vehi-cl e and launcher

1 camera alined onrange flight azimutt

to view fallingobjects

2 view5 shovingspacecra f t mbi l ica ldisconnect

3 view5 of space-c ra f t dur ing l i f t -of

2 view5 showing LESand CM

2 view5 showing LESsnd CM

Copies from VeroBeach, MelbourneBeach

s p i e s from VeroBeach, MelbourneBeach, Grand BahamaIsland, Cocoa Beach,3nd Patrick AFB

:opy from Cocoa Bead

CD - Calendar lay WD - Working Day

Reques ecSelivery

(a)

T+5 CD

T+5 0

T+ 5 CD

T+5 CD

T+5 CD

w

T+5

T+5

T+5

"+l CD

T+5 CD

T e CD

bCeta requested but no t received during the postlaunch reporting period.


Date/time

0

received

T+8 CD

T& CD

T+8 CD

(b 1

T+8 CD

T+12 CD

T + l 2 CD

(b1

(b

T+1 CD

T+l CD

T +8 CD

H - HOWS



7-30 U NCLASS I F I ED

20

52A

PI

l-0 ,

I

I ? I ' i





8

83

8W

0W

?

2

rl

In

7c'!rl

-

N

f

t

2

n Nl

:cu

:M:N

2(u

0:

M0

40NN

U N CLASS IF I ED



U N C L A S S I F I E D-3 2 II M 3

CXO

W P -1

N N

d rl

rl

-r ld

N

d

-

M

nl

N

6

'torl

N

d

U NCLASS I FI ED




r i d 0d c u

cuIn

P- v,

4 0 0r l c u

4-M

t Inri

5rne,

&

m




7-34

~

Apollo mating rin g, long itudi nal nio-801 s3-05003

Apollo mating ring, tang enti al El l- 80 1 s3-05002

Apollo mating ring, perpendicular EI2-801 s3-05001 I


Angle of at tack (pi tch , coarse) D133-900

Angle of attack (pit,ch, fine) D134-900

Angle of s ides l ip (yaw, coarse) D135-900

Angle of s id e sl ip (yaw, f i n e ) D136-900

m i c ressure (coarse D137-900

Dynamic pres sure ( f ine ) D138-900

TABLE 7. 4-IV. - LAUNCH-VEHICLE' DATA PROCESSED

FOR MSC-HOUSTON AT EFC

processed a t MSFC were requ ired between 7 and 12 calendaraf ter launch; dates of receipt are l i s t e d i n t a b l e 7.4-V 1

P1-B1-17

p i -~1-18

F6 x-1g-08

F6-x-19-09

P1-331-20

~6-x-19-10

Measurement

~011coarse)

Rol l ( f ine )

Yaw (coarse)

Yaw ( f ine )

Pi tch (coarse

Pi tch ( f ine )

hunch-vehicle

1124-802 F6-X-BU-08

H40-802 n-17M-03

II25-802 F6-X-BU-09

H41-802 I?j-14M-02

II26-802 F6-X-B12-10

H42-802 F6-14M-m

Pitch F42-802

Yaw F43-802

Roll F44-8O2

F6-08

~ 6 - 0 5

F6-04

U N CLASS IF1ED




I

?f

r-l

3+!3

aJ

Ln.-I+H

Mrl

% 2 &

(0

Q,4J

ldkacal

mQ,a?

4Jc,

4

x

to4J0r lPi

3





North

TowerComplex 37 B

Figure 7.4-2.- Engineering sequential fixed camera locations for Apollo m iss i on A-102.

U NCLASS I F I ED



7-38

-No. 2

B-10

B-11

B-12

B-14

B-16


No. 3

C-6

C-?

C-8

c-10

c-11

c-12

C-14

c-16

7.5 Telemetry Tape Se lec t ion and V er i f i c a t ion

Pr ior t o th e Apol lo r r. ission A - 1 0 2 , a procedure was e s t a b l i sh e d t oa s s u r e t h a t t h e best of the launch-s i te te lem etry tapes would be se lec ted

for processing the launch-phase telemetry da ta .

four launch-si te ground st .at ions were ob ta ined f o r the se l ec t ion p rocess .

They were from the Hangar S st at io n, Telemetry Building no. 2 (Tel 2) ,

Telemetry Building no. 3 (Te l 3 ) , and the Mission Control Center (MCC).

T3lemetry tapes from

A s soon as poss ib le a f t e r lo s s ot' s igna l , cop ies of the Hangar S

tape w e r e made a t the Hangar S playback s ta t ion , tape copies of t h e

Tel 2 and Tzl 3 t a p e s were made a t t h e Tel 2 s t a t io n and taken a t once

t o the Hangar S playback s ta t ion , and the o r ig ina l MCC tape w a s takent o t h e. H an g ar S playback s ta t ion .

c a rr i e rs were made from each of the se tap es w i t h t he iden t i za l f o rma t ,

galvonmneter de fl ec ti on , galvonometer frequency response , and disc rim-inator low-pass o u t p u t filters. The IRIG standard filters and galvan-

ometers were used. The format for the t h r ee o sc i l log raph r eco rd ings

w a s as follows:

Three oscil lographs of s e l e c t e d sub-

No. 1

A-6

A-10

A-11

A-12

A-13

A-14

A-15

A-16

The oscillographs made from the Hangar S t ape cop ies were comparedwith the near rea l - t ime oscillographs made from the Hangar S o r i g i n a l

tane t o v e r i f y t h e s e t a pe c op ie s.

The osci l lographs mad.e from the Hangar S, T ? l 2, and E l 3 tapecopies and from the MCC o r i g i n a l t a p e w e r e examined by a committee

cons i s t ing of representatives from MSC-Florida Ope ra ti on s and MSC-

Houston fo r overa l l s ig nal qual i ty , d ropouts , no ise conten t , and pos-

s i b l e s p i ke s of t h e t y p e e x hi b i te d i n t h e BP-13 d a t a ( r e f . 1 0 ) .committee judged the Hangar S and Tel 3 tape copies and the MCC o r i g i n a l

t ape t o be adequa te f o r the da ta r educ t ion p rocess.

The





The original MCC t ape w a s s e l e c t e d from t hese th r ee f o r d a t a pro-

cess ing for t h j s m i ss io n. The MCC tape covered a maximum time of recep-

t i o n a f t e r lift-off, and the 100-pps t iming s ignal on a s e p a r a t e t r a c kw a s , a t once, compatible with the processing equipment a t MSC-Houston.

The Te l 2 tape w a s t h e l eas t prefer red because of a ?-second dropout

which occurred approxim9tely 20 seconds before s tag ing . The committee 's

f inal judgment of the f o u r t a p es i n o r de r of preference was (1)Hangar S,

( 2 ) MCC, (3) T z 1 3, and ( 4 ) T e l 2.

U N CLASS IF1ED



UNCL AS S IF1ED

8 . 0 APPENDIX B

UNCLASSIF IED

8-1




.crtc a m

I N f l c [-

m*

ad

al

N

W

mf-

C

dc,

P

E

k

c,

a

W-

C

d

c)

P

+

Ed





W-o O FN & N N

O ~ W W W W

Id., m< P<P< P<x x x x x x

n n n n n n n i n mi , + d 4 r l r l r l + l r l

0 0 0 0 0 0 0 0 0o @ u * * , c r u u u

3 0 0 0 0 0 0 0 0

3 0 0 0 0 0 0 0 0i d r l d r l r i d d r l

$Lo9)

a




8-4

I

I

6

U NCLASS IF1ED

2 2 2 2 2 2 2 2 2 2 2 2 20 0 0 0 0 0 0 0 0 0 0 0 0* * Y * Q Q Y Y Y Y U Y U

0 0 0 0 0 0 0 0 0 0 0 0 0

U N CLASS I F I E



U NCLASS F IED

n

E F F h

n w - m

k k k k

8-5




U N C L A S S I Fl E D 8-7

U NCLASS IFI ED



U NCLASS IF l ED-a

I---

X

D

kX WO EO8

.d -$ 23

+-O U U OI I I I L Ca a 5 a

? ? . > B

lr.l

J E

U NCLASS I F I ED



r.

UNCLASS F IED

H I 1 I I

a a o 0 o o o o ac, c , u c , c , w I

m a 0 0 0 0 0 0 ' ;

41 I

V

I I

8-9





I----

S S S B S0 0 0 0 0




TABLE 8.2-1.- PERATIONAL TEST PROCEDURES FOR APGLLO SP-15

SPACECRAFT AT CONTRACTOR'S MANUFACTURING FACILITY

OTP

UNCLASS IF IED

P-3022

P- 018

P-8019

P-1040

P- 019

P-3036

P-3033

P-3000

P-10002

P-1008

P-E169

P-8169

P-1095-1094

P-3015

P-3071

P-3013

P-8077

P- 003

P-5000

P-3014

P-3072

P-3015

T i t l e

LES horizontal weight and balance

Forwayd hea t -shie ld im ta l l a t io n

Beacon systems check-out

U S heck-out

GE E integrated check-out

CM v er t i ca l weight and balance

CM horizontal weight and bahmce

ECS f i l l and check-out

Test conf igura t i on check l i s t

El e ct r ic a l power systems check-out

Telemetry and instr umen tatio n check-out

(Recycle )

Bat te ry seFV c ng

Adapter and SM s tack and a l ine

CM s tack and a l ine

LES e tack and a l ine

Omnibeacon system check-out

Integra ted sys tems ' t e s t

ECS drain and purge

Demate LES

Demate CM

Demate SM

8-11

Dace performed

(1964)

Mar. 12 and 13

A p r . 7 t o 13

Apr. 23 t o 24

May 1L t o 18

Apr. 21 t o 24

Apr. 20 t o 22

Apr. 22 and 23

May 7 t o May 22

Apr. 23 t o May 19

Apr. 27 t o 30

hpr . 30 t o May 5

May 8 t o 1 4

Apr. 29 t o May 28

May 11 t o 13

May 1-3

Mzy 14 and 15

Apr. 23 and 24

yay 18 t o 25

26

MELY 28

YAY 29

May 27 and 28

U N CLASS IF1ED



8- 2 U N C L A S S I F I E D

TABLE 8.2-11,-OPERATIONAL TEST PROCEDURES FOR A P O U O BP-15

SPACECRAFT AT FLORIDA OPEXATIONS

O T P

C-0004

c-0005

c-0006

c-0007

c-0009

c-0021

C-0028

c-0031

(2-0033

c-1012

C-3044

C-3045

c-3063

c-3065

Title

-~

Electrical interface checks

Integrated systems check-out withlaunch-vehicle simulator

Spacecraft-launch-vehicle overall

test no. 1 (plugs in)

Spacecraft-launch-vehicle countdown

Spacecraft-launch-vehicle RFI test

Spacecraft-launch-vehicle overalltest no, 2 with ordnance in andplugs ou t using swing arms

Spacecraft-launch-vehicle simulatedflight test

Launch-vehicle sequencer malfunction -spacecraft monitor test

Spacecraft-launch-vehicle countdown

demonstration

Battery charge, discharge, all batteries

U S total weight and center-of-gravity

determination

Buildup and assembly of LES

Spacecraft off-loading, transportation to

Hangar AF, and preparation for receivinginspection

Transportation of spacecraft to launchcomplex and mating of spacecraft to

launch vehicle

Date performed

(1964

August 7

July 9

Aug- 19

Sept. 17 and 18

Aug . 24

Aw. 29

Sept. 3

A u g . I 2

Sept. 14 and 15

Sept. 16

May 21

July 1 to Aug . 3

June 8 o 15

June 26

U NCLASS IF1ED




T A B U 8.2-11.- OPERATIONAL TEST PROCEDURE FOR APOLLO BP-15

SPACECWT AT FLORIDA OPEEUTIONS - Concluded

OTP

c-8112

C-8114A

c- 8115

c-8131

c-8132

c- 9002

c-go36

C- 9037

c-1oool.A

T i t l e

Instrument system PIA procedures

Antenna (VSwIi, t e s t )

RF systems PIA procedure

Telemetry systems t e s t

Power and s eq u en ti al PLA pro ced ure

(tower sequencer)

F u n c t i o n a l v e r i f i c a t i o n of A 1 4 - 0 0 1

launch-escape tower s u b s t i t u t e u n i t

W a t e r -g ly ol m i c he ck- out

Ground equipment Ve r i f i ca t io n and GSE

integrated umbil ical check-out

Pad checkl i s t

~~ ~

Date performed

(1964

Continuing

June 29

Continuing

July 1 and 2

J u l y 23

and 31Sept. 2, and 10

June 9

AX. i o , 21, 30,

June 23

June 15

aPerformed i n suppor t of all launch-vehicle-spacecraft

i n t egra t ed t e s t s .




U N CLASS IF1ED

9.0 APPENDIX c

9.1 C-Band Beacon Anomaly

9-1

Seven tracking network radar stat ions reported seeing two beaconsdu r ing the BP-15 s p a c e cr a f t 's f i r s t o r b i t a l p as s ove r t h e i r s t a t i on s .

Antigua Is la nd and Ascension Isl and r eport ed de te ct in g th e two beacons

on t he spacec ra f t ' s second o r b i t a l pass. Thi s two-beacon resp onse wasnot rep orted on previous Saturn launches; consequently, th e repo rt w a s

unexpected and re su lt ed i n some operato r confusion during tr ack ing

ope r a t ons.

In a l l cases , th e s t a t i ons de tec t ing the two beacons t racked onth e beacon I t . . . fa r t he st out in range." However, i f t he ope ra to r

had e l ec ted t o t r ac k the "c lo ses t " beacon, t r ack ing e r r o r s of 500 yardswould have been introduced on the f i r s t two orb i ta l passes.

The instrument unit ( I U ) beacon transponder w a s connected t o the

sho r t - l i f e b a t t e ry bus ( l eng th of ope ra t ion - approximately 40 minutes)on previous Saturn launches , wi th th e re su l t th a t the t ransponder went

i n op e ra t iv e p r i o r t o t h e f i r s t o r b i t a l pass over Ascension Island. On

SA-7, t h e I U beacon was connected to the long - l i f e ba t t e ry bus which

r e s u l t e d i n a probable operation of approximately 110 minutes.

For t ra ck in g purposes , the BP-15/SA-7 pace veh ic le u t i l i zed twoC-band beacon transpo nder s: one i n t h e I U , and th e o the r i n the space-

craf t opera t ing on ad jacent f requencies ( l e s s t h a n 1mc d i f f er e n c e ).

Two sep ara te redundant beacons operat ing on th e same spacec ra f t f r e -quency w e r e installed within t h e command.module. The beacons were

interrogated by ground-tracking radars u t i l i z i n g a double pulse in ter-xga t ion sys tem as shown in the following diagram.

Pulse w i d t h = 1 psec-

1 1

- determined by PRF - -

j ( 4

I(1538 Ksec o r gr ea te r)





Pulse ( a ) t r i g g e r s t h e I U beacon which returns response pulse ( A )

2s shown below. Pulses (a) and ( b ) t r igge r the spacec ra f t t r ansponder

xhich re turns response pu lse (B). Although the I U and spacecraf t t r a n s -ponders operate a t s l i g h t l y d i f f e r e n t f r eq u e nc i es , b ot h a re wi t h i n t h e

ground radar bandwidth. The 3.5-u,sec delay appears t o t he ope ra to r a st,wo t a r g e t s 500 y w d s a p ar t .

T - b l e 9.1-1 i s a summary of s e l e c t ed radar r e p o r t s , w i t h Ascension

I s l and be ing the l a s t s t a t i o n t o see the two beacons a t l : l 7 : 56 e. s. t.S u b t r a c t i n g t h e l i f t - o f f t i m e of ll:22:43 e. s. t . from the Ascensiontime.shows t h a t t he beacon became inopera' ble sometime a f t e r 115 minutes

a f t e r l i f t - o f f . T h i s time compares fa vor abl y with th e MSFC pre l iminaryes t ima te for opera t ion of approximately 133 minutes.

The t ransponder t r igger ra te was t e l eme te red t o g round s t a t io ns

and i s shown i n f ig ur e 9.1-1.

frequency (PRF) r a t e s d x i n g t h e p er io ds of beacon interrogation.

r e l a t i o n o f t h e se data ind i ca t e proper opera t ion of th e t ransponder .

Also shown a re t h e radar p u l s e r e p e t i t i o n

Cor-

N o a d d i t i o n a l i n v e s t i g a t i o n o f t h i s a no na ly i s plmned.





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F igure 9.1-1.- B P -1 5 spacecraf t t ransponder in terrogation and response.





9.2 Loss of Temperature Measurement ~ ~ 5 8 7 7 ~

9-5

Temperature data from thermocouple SR5877T fa i l e d t o in di ca te an ychange during the BP-15 spacecraf t mission.

ca ted 1 inch from the r i m of the nozzle on the pos i t i ve p i t ch engine on

th e in stru men ted RCS quad (fi gs . 4.8-7 t o 4.8-9). The temperature

measurement indicated a n output reading of approximately 4 ercent off u l l sca le , equ iva len t t o 104' F, and remined constan t dur ing count-

down and f l igh t .no t agre e wi th th e expected environment ind ic a ted by o the r re la te d

thermocouple measurements on the RCS quad dur ing f l igh t .

The thermocouple was 10-

This val ue coul d be pos si bl e durin g countdown but di d

Data from t h e thermocouple w a s highl y des i ra b le because t he temper-

ature of t h i s par t icular area was predic ted t o exceed a l l ot he r temper-a ture measurements on th e RCS quad. Also, temperature data f rom th is

area could have been used i n the inve st iga t ion of tne l o s s of data fromca lo r ime te r no. 13 loc a te d immedia te ly below t h i s nozzle ( f ig . 4.11-2).

The instrumented RC S quad w a s in st al le d on th e se rv ic e module on

th e la unch pad on June 29 , 1964. A t e l eme t ry t e s t was performed on

July 2, 1964, duri ng which a l l thermocouples were thermally excited

and t e s ted us ing a heat gun. All th e RCS ins trum enta tion fu ncti oned

s a t i s f a c t o r i l y a t th e time. The same measurement w a s e l e c t r i c a l l y

c a l i b r a t e d 2 minutes before launch and d i d not indicate any anomaly a tt h e t i m e .

A n a l y s i s o f f l i g h t data d id no t ind ica te any s ol ut io n t o t h i s

anomaly. The results o f a d d i t i o n a l s t u d i e s w i l l be reported when avai l -able .





9 .3 L o s s of Heat Flux Data from

Calor imeter 13 (SA0353R)

Heat flux data f rom calor imeter 13 (SA0553R) was quest ionable and

hed t o be d isregarded .

nozzle of the po si t i ve p i tc h engine of the ins t rumented RC S quad

( f i g . 4.11-2). The ca lor im et er gave normal resp onse when the rm all y

exc i t ed by a he at gun during th e s im ul at ed f l i g h t t e s t a t T-15 days and

aga in r esponded no r rr a lly du ring e l ec t r i c a l ca l ib r a t i on a t approximately

T -2 minutes. The cal ori me te r d id r e spond to exc i t a t ion a t l i f t - o f f .

I t fai led, however , t o respond a t t h e w i n h e a t p u l s e d ur i ng a s c e nt ,

whereas the body temperature measurement SA0563T located inside the .

same cal or im eter in crea sed to a l e v e l above t h a t o f t h e o t h e r c a l o r i -

meter body temperatures. This indi cate d t ha t a n unmeasured heat f luxwas presen t i n t h i s pa r t i cu l a r a r e a of the se rv ice module.

deta d i d not g ive any informat ion with respect t o th e t i m e or the causeof t h e failure.

The calorimeter w a s l o c a t e d on t he SM under the

T h e f l i g h t

The data from t h i s c alor imet er were highl y des i rab le because th i sa r ea w a s pred ic ted t o have t h e h ighes t hea t ing r a t e on the se rv i ce

module.s p a c e c r a f t a l s o f a i l e d .

f ac to ry da ta on the BP-13 s p a c e c r a f t f l i g h t , b u t pe rf or me d s a t i s f a c t o r i l y

on the BP-15 s p a c e c r a f t f l i g h t .

The same type of ca lo r ime te r i n the same location on the BP-13Two o the r ca lo r ime te r s f a i l e d t o produce sat i s -

Calorimeters of two di ff er en t ranges were used i n bo th f l igh t s .

2The 0 t o 25 Btu / f t / see ca l o r i ne t e r s loca te d on th e c o m n d module pe r -

f x m ed s a t i s f a c t o r i l y i n bo th f l i g h t s .meters, l oca ted on th e SM, dif fered f rom the ca lor imeter loca ted on

the CM i n si ze . The diaphragm of th e CM ca lo r ime te r w a s 3 .5 m i l s t h i c k

x i t h 0.170 inch unsupported diameter and t ha t of the SM ca lo r ime te r was2 .5 mils th i ck wi th 0.250 inch unsupported diameter ( f i g . 9-3-1).

The 0 t o 5 B t u / f t 2 / s e c c a l o r i -

A s a r e s u l t o f t h e f a i l u r e s du r in g t h e BP-13 s p a c e c r a f t f l i g h t ,

PGC-lESD per formed addi t ional environmenta l tes ts on the smal ler range

ca lo r ime te r p r io r t o the BP-15 spacec ra f t f l i gh t .which t he cal ori me te r specimens were sub jec ted were based on th e BP-13

spac ecra f t launch environment. The t ime dura tion during aco ust ic nois e

t e s t , however, exceeded th e launc h environment. None of th e environ-

men ta l t e s t cond i t ions were imposed simul taneous ly. The followi ng a r e

the environmental condit i ,ons t o which the calo r ime ter w a s subjec ted :

The t e s t l eve l s unde r

UNCLASS F l ED



9-7

a.

b. Vibration

Humidity - 100 percent a t 90' F t o 150" F f o r 18 hours

(1) Pa ra l l e l t o the diaphragm

0.32-inch D. A. D.

lOOg peak-to-peak

120g peak-to-peak

i o t o 60 CPS

60 t o 1,000 CPS

1,000 t o 2,020 cps

lOOg peak-to-peak 300 cps for 15 seconds

( 2 ) Perpendicular t o t he diaphragm

0.32-inch D. A. D. io t o 60 CPS

120g peak-to-peak 60 t o 2,020 cPs

120g peak-to-peak 300 cps for 15 seconds

Vibrat ion cycle time i n each di rec t io n was approximately

5 minutes and 1 5 seconds

e. Acoustic no ise l eve l s

Frequency Level Time

60 t o 2,400 cps 152 db 15 seconds

60 t o 2 ,400 cps . 159 db 30 seconds

60 t o 4,300 cps i6 6 db 3 minutes

6 db/octave roll-off on high s ide

d. Thermal shock was accomplished by a l t e rn a t e ly hea t i ng t o f u l l

2utput of sensor then cool ing wi th CO spray.

The ca lor imeters passed a l l t e s t s and a malfunct ion could not bedupl i cat ed i n t he l abora to ry .

The acous t i c no i se l eve l t e s t s were performed a t Space and Informa-See reference 11i o n Systems Div isio n of North American Avia tion , Inc .

for t he t e s t desc r ip t i on and re su l t s .



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Post Launch Report for Apollo Mission a-102

Documents

Transcript of Post Launch Report for Apollo Mission a-102