ASTP [SA-210) LAUNCH VEHICLE OPERATIONAL … FLIGHT TRAJECTORY DISPERSION ANALY SlS ... Launch...
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Contract WAS8-4016 Apollo/Soyuz Test Project ASTP [SA-210) LAUNCH VEHICLE OPERATIONAL FLIGHT TRAJECTORY DISPERSION ANALY SlS VOLUME I https://ntrs.nasa.gov/search.jsp?R=19760004113 2018-07-10T05:27:32+00:00Z
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Transcript of ASTP [SA-210) LAUNCH VEHICLE OPERATIONAL … FLIGHT TRAJECTORY DISPERSION ANALY SlS ... Launch...
Contract WAS8-4016 Apollo/Soyuz Test Project
ASTP [SA-210) LAUNCH VEHICLE OPERATIONAL FLIGHT TRAJECTORY
DISPERSION ANALY SlS
VOLUME I
https://ntrs.nasa.gov/search.jsp?R=19760004113 2018-07-10T05:27:32+00:00Z
i I
TECHNICAL NOTE
DRL 444-V4a
ASTP (SA-210) LAUNCH VEHIrLE OPERATIONAL FLIGKT TRAJECTORY
DISPERSION ANALYSIS
VOLUME I
A p r i l 4 , 1975
CONTRACT NAS 8-4016
APOLLO/SOYUZ TEST PROJECT
FLIGHT TECHNOLOGY BRANCH
CHRYSLER CORPORATION SPACE DIVISION
Prepared by: N . W. Will iams, G. W. Klug and F. A. Ransom
Approved by:
Performance and Mission Analysis Group
,ED. 3+ R. D. Taylor , 'Managing Engineer F l i g h t ~ e c h a n i c s s e c t i o n -
& Swider, Manager Technology Branch
' J ,,d &/dL//- .
R. H. Ross, Deputy P r o j e c t Manager Vehicle Systems
FOREWORD
This document is Data Requirements Lis t (DRL) Item 444-V4a. It is
Volume I of t he two volume documentation required f b r the ASTP (SA-210)
Launch Vehicle Operational F l igh t Trajectory Dispersion Analysis under
Contract NAS 8-4016, Schedule 11, Modification MSFC-1, Amendment 199. The
associated Guidance Hardware Error Analysis i s documented separately, A S
Volume 11, because it contains c l a s s i f i e d mater ial .
Acknowledgements a r e made t o the personnel of Marshall Space F l ight
Center SA&I-EL24 f o r t h e i r ass i s tance and cooperation.
TABLE OF CONTENTS
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FOREWORD
TABLE OF CONTENTS . . . . . . . . . . . . . . . . . . . . . . . . . . LIST OF TABLES . . . . . . . . . . . . . . . . . . . . . . . . . . . DEFINITIONS AND SYMBOLS . . . . . . . . . . . . . . . . . . . . . . . SUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 . 0 INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . 2.0 DISCUSSION . . . . . . . . . . . . . . . . . . . . . . . . . . .
2.1 M i s s i o n D e s c r i p t i o n . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.2 Launch Veh ic l e and T r a j e c t o r y D e s c r i p t i o n
2.2.1 L a u n c h v e h i c l e . . . . . . . . . . . . . . . . . . . 2.2.2 F l i g h t Environment . . . . . . . . . . . . . . . . .
. . . . . . . . . . . . . 2.2.3 F l i g h t Sequence of Events
. . . . . . . . . . . 2.3 Di spe r s ion E r r o r Sources
2.4 T r a j e c t o r y Di spe r s ion and A n a l y t i c a l Procedures
3.0 RESULTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.1 T r a j e c t o r y Di spe r s ions
. . . . 3.2 S-IVB S t a g e F l i g h t Performance Reserve
. . . . . . . . . 4.0 GOVERNMENT FURNISHED DOCUMENTATION
5.0 REFERENCES . . . . . . . . . . . . . . . . . . . . . DISTRIBUTION . . . . . . . . . . . . . . . . . . . . . .
Pag_e
i
i i
iii
iv
1
4
5
5
5
5
6
6
7
8
Table - 1
2
3
4
LIST OF TABLES
Vehic le Weight Breakdown . . . . . . . . . . . . . . . . . . . . 16
. . . . . . . . . . . . . . . . . . . F l i g h t Sequence of Events 1 7
Three Sigma Tolerances . . . . . . . . . . . . . . . . . . . . . 18
T r a j e c t o r y Dispers ions a t S-IB/S-Ivs Separa t ion , S-IB Propul ;ion/Non-Propulsion Three Sigma Deviations . . . . . . . . 20
T r a j e c t o r y Dispersions a t S-IB/S-IVB Separa t ion , S-IVB . . . . . . . . Propulsion/Non-Propulsion Three Sigma Deviations 25
T r a j e c t o r y Dispersion Envelope a t S-IBIS-IVB Separa t ion , Combined S-IB and S-IVB Stage Three Sigma Deviations . . . . . . 3 0
T r a j e c t o r y Dispersions a t Orbi t I n s e r t i o n , S-IB . . . . . . . . Propulsion/Non-Propulsion Three Sigma Deviations 35
T r a j e c t o r y Dispersions a t Orb i t I n s e r t i o n , S-IVB . . . . . . . . Propulsion/~on-Tropulsion Three Sigma Deviations 43
T r a j e c t o r y Dispers ion Envelope a t Orbi t I n s e r t i o n , Combined . . . . . . . . S-IB, S-IVB Stage and IMUThree Sigma Deviations 51
Three Sigma F l i g h t Envelope of P e r t i n e n t Design Parameters, . . . . . . . . . . . . . . . . . . . . . . . F i r s t Stage F l i g h t 58
. . . . . . . . Performance Trade-offs a t S-IB/S-IVB Separa t ion 60
. . . . . . . . . . . Performance Trade-offs a t Orbi t I n s e r t i o n 61
Large Guidance Pla t form Azimuth Misalignment E f f e c t s a t 6 2 O r b i t I n s e r t i o n . . . . . . . . . . . . . . . . . . . . . . . .
. . . . . . . . . . . . . S-IVB S tage F l i g h t Perf orrnance Reserve 63
DEFINITIONS AND SYMBOLS
Aerodynamic Heating Indicator qVr d t q ;- dynafiic pressure V r : relatSve velocity
1 I a t = tota! :angle of at tack
Aerodynamic Load Indicator Product of dynamic press.,re 3nd sngle of a t tack .
Altitude
Angle of Attack, Pitch
Apogee Altitude
Attitude Command
Attitude Error
Axial Force
Descending Node Argumsnt
Dynsmic Pressure
Vehicle a l t i t u d e above the reference e l l ipsoid measured aiong the geocentric position vector.
Angle between the pi tch ?lane compocent of the r e l a t i v e velocity vector and the ~ o n g i - tudinal axis of the vehicle, measured poaitive nose up.
Apogee height of the osculating conic above the reference e l l ipso id , referenced t o the equatorial radius, 6378165 m t e r s . Eulerian angle command,derived by the guidance system and transmitted t o the control system.
Difference between the vehicle a t t i t u d e (p i tch , yaw and r o l l Eulerian angles) and the vehicie a t t i t u d e command.
Component of the r e su l t an t aerodynamic force along the vehicle longitudinal axis (X axis of PASCS h), measured posi t ive toward the nose of the vehicle.
Angle meaaured in the equatorial plane between t h e o rb i t descending node and the space fixed launch meridian defined a t Guidance Reference Release.
Component of the r e su l t an t aerodynamic force along the relatLve velocity vector. meerured pos i t ive opposite t o the veloci t vector.
4 x ( ~ s i t y ) x (Relative
KBFINITIONS AND SYMBOLS (CONT ' D)
Earbh Fixed Cross Range Ye component of PUCS 10 posi t ion vector .
Ear th Fixed Fl ight Path Angle Angle between the ea r th fixed ve loc i ty vector and the ea r th fuced geocentric pas i t ion vector (PASCS 11 ) , measured pos i t ive downrange from the pos i t ion vector.
Earth Fixed Posi t ion Posi t ion vector/cornponents i n an earth-fuced pad-centered plumbline coor.. -.;late system. The Xe ax i s i s co inc ider~t with t k e reference e l l i p so id normal, pos i t ive upward. The Ze a x i s i s p a r a l l e l to the earth-fixed ainling azimuth and is pos i t ive downrange. The Ye d s com- p l e t e s a r i g h t handed sys tcn . ( p ~ S l . 5 10)
Ear th Fixed Velocity Velocity vector/components i n FASCS 10.
i e 2 Z L
Ear th Fixed Velocity Magnitude + f e i " Le
Eccen t r i c i t y Eccent r ic i ty of t he osculat ing conic.
F l igh t Azimuth
Geocentric Declination
Geodetic Lat i tude
Ground Range
I n c l i n a t i o n
I n e r t i a l Range Angle
Angle def in ing or ien ta t ion of t he space f i xed coordinate system downrange ax is , Zs, a t Guidance Reference Release, measured pos i t i ve e a s t of nor th i n plane normal t o t he space f ixed X s axls.
Angle between the geocentr ic rad ius vector and the t r u e equa to r i a l plane, measured pos i t i ve north of the equator.
Angle between the reference e l l i p s o i d normal through t h e point of i n t e r e s t and the t r u e equa to r i a l plane, measured pos i t i ve north of the equator.
Surface dis tance from launch s i t e t o t he sub- vehicle point , pos i t ive e a s t (0" - 1800).
Angle between the instantaneous f l i g h t plane and the equa to r i a l plane.
Angle between the instantaneous space f ixed pos i t ion vector and the space fixed pos i t ion vector a t Guidance Reference Release.
Longitude
DEFINITIONS AND SYMBOLS (COW ' D)
Angle between the Greenwich meridian plane and the projection of the geocentric poei- t i o n vector i n the equatorial plane, measured posi t ive eas t of Greenwich.
Longitudinal Acceleration That par t of the t o t a l measurable acce lera t ion directed along the longitudinal ax i s of the vehicle.
Mach Number (Relat ive Velocity) f (Local Speed of Sound)
Ma88 Mass of the vehicle .
Navigation Coordinate System This system is ident ica l t o PASCS 13 with idea l navigation.
Normal Force
Perigee Alt i tude
Magnitude of t he resu l tan t aerodynamic fo rce normal t o the vehicle longitudinal ax i s , and i n t he plane defined by t h a t ax i s and the r e l a t i v e veloci ty vector.
Perigee height of the osculat ing conic above the reference e l l i p so id , referenced t o t h e equatorial radius , 6378165 meters,
Period Period of the osculat ing conic.
Pi tch, Yaw, Roll ( I n e r t i a l ) Eulerian angles of vehicle a t t i t u d e measured with respect t o the space fixed coordinate system. Vehicle a t t i t u d e is defined by t h e ordered ro ta t ion of p i tch , yaw, and r o l l . (See i l l u s t r a t i o n )
Radius Space fixed poeition vector magnitude,
Range
Relative Vehicle At t i tude
Surface dis tance from launch r i t e t o t he eub-vehicle point , poai t ive e a r t (0' - 180').
Pi tch, yaw and r o l l angles of t h e vehic le i n a n ea r th r e l a t i v e system. The r o l l urir ir t h e project ion of the ve loc i ty vector i n the loca l horizontal plane; the yaw ax i r i e i n t h e loca l v e r t i c a l plane, pos i t ive toward the center of the ear th ; the pi tch a x i s complrtw a r igh t handed syetem. Vehicle a t t i t u d e lm defined by ordered rotat ion-pi tch, yaw and r o l l .
Launch Mecidkn
Pitch Yaw Rol I
Relative Velocity
Semi-Ma j o t Axis
Velocity r e l a t i v e t o the atmoephere ( in- cludee wind ve loc i ty) .
Length of the chord in the o r b i t plane connecting t h e apogee and the perigee of the osculat ing conic.
Space Fixed Croee Range Ya component of PASCS 13 pos i t ion vector .
Space Fixed Fl ight Path Angle Angle between the apace fixed veloci ty vector and the radius vector (PASCS 13), measured poei t ive downrange from radius vector.
Space Fixed Position Poaltlon vector/componente i n a space fixed ear th centered, plumbline coordinate system defined a t Guide-ce Reference Releare. The Xe axle i a p a r a l l e l t o the reference e l l ipeoid norm1 which paesee througb t h e launch s i t e . The 2s axis i s p a r a l l e l t o , and poei t ive i n the same d i r ec t ion as, t h e earth-f ixed f i r i n g azimuth. The Y s u l r cmple tee the r igh t handed aystem. Thir is Project Apollo Standard Coordinate System 13. (PASCS 13).
Space Fixed Velocity Velocity vector/components in PASCS 13.
Space Fixed Velocity Magnitude T
28
T ime Inst8ntaneous f l i g h t time referenced t o f irst mu: '22.
Three Sigma (36) Three standard deviations.
Thrus t
True Anomaly
Velocity Vector Azimuth
Tota l e f f ec t ive th rus t magnitude,
Angular dieplacement of t he vehicle C.C. from t h e perigee, measured i n the d i r ec t ion of the motion.
The angle between t h e ve loc i ty vec tor pro- jec t ion on the ear th ' s au r f r ce and truo north.
v i i i
Vehicle Weight
Y 4 Posi t ion Vector
Inetantaneoue t o t a l vehicle weight.
Vehicle c. 8. dieplacement component8 in a apace f ixed, r i gh t handed, t a r g e t coordinate ayetem with i t 8 o r ig in a t t h e center of t he ear th. The Xq a x i r parser through the descending node of t he orbit plane. The 24 axis l i e 8 i n t h e desired o r b i t plane 90' downrange from the & axis . The Y4 ax i s completes a r igh t handed eyetem and i e perpendicular t o t h e orbit plane.
AFETR
AFB
AH1
APS
APSO
AS
ASTP
B- 7
C
c/o
CS
CCSD
C .G.
CM
CSM
DELTA ( )
(AP)
DIA
DM
DRL
ECF
EMR
F
F PR
FT
8
DEFINITION AND SYMBOLS (CONT ' D)
A i r Force Eastern Test Range
A i r Force Base
Aerodynamic Heating lnd i cu to r
Auxiliary Propulsion S y s t ~ n , ~ ~
Apollo Soyuz Program O f f i c c
Apollo Saturn
Apollo Soyuz Test Project
Trajectory Data Tape
C-Band Radar S t a t ions
Cut off
Command System
Chrysler Corporation Space D i ~ i s i o n
Center of Gravity
Command Moduie
Command and Service Modules
Increment
Parameter Incement
Vehicle Diameter
Docking Module
Data Requirements L i s t
End Conditions of F l igh t
Engine Mixture Ratio
Average Longitudinal Sea Level Thrust
F l igh t Performence Reserve
F l igh t Technology
Acceleration of Gravity a t Sea L w e i (9.80665 rn/sec*)
GCS
GET
GRR
GSFC
H- 1
I B M
IECO
IGM
LAC
LC
LES
LH2
LMS C / HREC
LOX
LSA
LVDC
DEFINITIONS AND SVME3LS ( CONT ' D)
Guldance Cutoff S igna l
Ground Elapsed Time
Government Furnished Documentation
Gaseous Hydrogen
Greenwich Mean "me
Guidance Reference Release
Goddard Space F l i g h t Center
S-IB Stage Engine
I n c l i n a t i o n
I n t e r n a t i o n a l Business Machines Corp.
Inboard Engine Cutoff S igna l
I t e r a t i v e Guidance Mode
I n e r t i a Measurement Unit
S p e c i f i c Impulsc?
Instrument Unit
S-IVB Stage Enginp
Kennedy Spacef 1 igh t Center
Loss of A t t l t u d e Control
Launch Complex
Launch Escape S y ~ t e m
Liquid Hydrogen
Lockheed M i s s i l e s and Space Company/Huntsville Research and Engineering Center
Liquid Oxygen
Level Sensor Actuation
Launch Vehicle D i g i t a l Computer
L/V
LWC
LWO
MDAC
MSFC
NASA
NPV
OECO
PASCS
POT
PSF ; psf
9
RP- 1
RS S
S-IB
S-IU
DEFINITIONS AND SYMBOLS (CONT ' D)
Launch Vehicle
Launch Wi~dow Closing
Launch Window Opening
McDonnell Douglas A i r c r a f t Corporation
Marshall Space F l i g h t Center
National Aeronautics and Space Adminis t ra t ion
Liquid Nitrogen
Not Applicable
Non-Dimens iona 1
Non-Propulsive Vent
Outboard Engine Cutoff S igna l
Orbi t I n s e r t ion
Operat iona 1 T r a j e c t o r y
Pro jec t Apollo Standard Coordinate System
Prel iminary Operat ional F l i g h t T r a j e c t o r y
Pounds Per Square Foot
Dynamic Pressure
S-IB Prope l l an t
Root-Sum-Square
F i r s t Stage of t h e Saturn I B Launch Vehic le
Saturn Uunch Vehicle Instrument Vehicle
Second Stage of t h e Saturn I B Launch Vehic le
Saturn
Spacecraf t
x i i
SIGMA ( o )
( 2 )
SLA
SM
ST A
T
UHF
USA
VHF
DEFINITIONS AND SYMBOLS (CONT 'D)
Standard Deviation
Summation of
Spacecraft Launch Adapter
Service Module
Vehicle S ta t ion Location
Telemetry S ta t ions ; Time Base One Time
Time Base Zero
Time Base One
Time Base Two
Time Base Three
lime Base Four
Technical Note
Ultra High Frequency
United S ta tes of America
Union of Soviet Soc ia l i s t Republics
Very High Frequency
Flowrate
x i i i
SUMMARY
This r e p o r t p resen t s t h e ASTP (SA-210) Launch Vehic le Operatiorla1
F l i g h t T r a j e c t o r y (OT) t h r e e sigma ( 3 r ) f l i g h t parameter envelopes, t h e
S-IVB Stage F l i g h t Performance Reserve (FPR) , t h e S-IB s t a g e design parameter
envelopes, and p e r t i n e n t t r ade -of f f a c t o r s . The ASTP (SA-210) Launch Vehicle
500 Pound Launch Window Opening Operat ional F l i g h t T r a j e c t o r y was u t i l i z e d
a s t h e nominal f o r t h i s a n a l y s i s .
The f l i g h t envelopes presented a r e t h e r e s u l t s of s t a t i s t i c a l com-
b i n a t i o n s of p e r t u r b a t i o n e f f e c t s , eniploying t h e Root-Sum-Square (RSS)
technique. Concise summaries of p e r t i n e n t t r a j e c t o r y parameter d i s p e r s i o n s
a t S-IBIS-IVB Separat ion and Orb i t I n s e r t i o n (01) fo l low.
F l i g h t Time ( sec )
Radius (m)
S-IBIS-IVB Sep. Orbi t I n s e r t i o n
+RS S - -RSS - +RS S - -RSS - 2.82 2.65 10.99 10.46
2050. 2303. 505. 502.
Space Fixed Ve loc i ty (m/sec) 46.10 41.82 2.47 2.46
Space Fixed Path Angle (deg) 1.872 1.776 0.018 0.018
Ground Range (m) 4646. 3607. 40192. 38336.
Ear th Fixed Cross Range (m) 4002. 2377. 4928. 5125.
I n c l i n a t i o n (deg) ----- ----- 0.019 0.019
Descending Node Argument (deg) ----- ----- 0.019 0.019
Three sigma ( S I T ) v a r i a t i o n s i n t h e es tab l i shments of t h e Launch
Vehic le D i g i t a l Computer (LVDC) t ime bases T2, T 3 , and T4 a r e d isplayed i n
t h e fo l lowing t a b l e . Also included a r e t h e 3 0 v a r i a t i o n s i n t h e time t h a t
dynamic p ressure (q) decreases t o one pound per square foo t ( p s f ) . The
time bases i n i t i a t e independent event sequences and q 5 1 psf is a
primary Launch Escape System (LES) j e t t i s o n i n g c r i t e r i o n .
T3 'Lime of T2 T3 T4 q = 1 psf
(sec ) ( s e ) (sec) ( s e c )
RSS (+) 2.70 2.82 10.99 3.37
RSS (-) 2.53 2.65 10.46 4.83
The 36 dev ia t ions i n 5-2 engine i g n i t i o n and Engine Mixture Ra t io (EMR)
s h i f t times a r e t h e same a s those shown f o r T3, s i n c e they a r e programmed
T3 even t s .
The S- IVB s t a g e t h r e e sigma F l i g h t Performance Reserve (FPR) r e q u i r e -
ments f o r t h i s launch a r e 1172 pounds of LOX and 683 pounds of LH2. Th i s
FPR i s considered t o be v a l i d a t any point i n t h e prescr ibed 500 pound launch
window, s i n c e a previous a n a l y s i s has e s t a b l i s h e d t h a t FPR v a r i a t i o n w i t h i n a
l a r g e r 700 pound launch window is l e s s than 50 pounds. U t i l i z i n g t h e s e FPR
d a t a , t h e t a b l e on t h e fo l lowing page provides an assessment of t h e S-IVB
r e s i d u a l p rope l l an t s p red ic ted f o r t h e nominal miss ion. The nominal launch
time i s 2.84 minutes p r i o r t o t h e p lanar f l i g h t oppor tun i ty , consequently,
96 pounds of t h e 500 pound launch window p r o p e l l a n t a l l o c a t i o n a r e r equ i red
f o r yaw s t e e r i n g t o t h e prescr ibed t a r g e t cond i t ions .
Tot ; l on board a t GCS
(1);;nutscable
(2)Total Available
LOX LH2 Total (Pounds) (Pounds) (Pounds)
3 s i :ma FPR a l loca t ion 1172 683 1855
Remc 1.ning launch window a l loca t ion 334 - 70 - 404 (4 .8 .1 EMR)
- Tota l a l l oca t ion 1506 7 53 22 59
Excess ava i lab le over a l l oca t ion 2 62 9 1 3 53
Excess useable a t 4 .8: l EMR 2 62 - 5 5 - 317 - Excess bias 0 36 36
(1) Unuseable determined by MSFCIMDAC t o assure the required 6.7 m/sec deplet ion cutoff t h r u s t decay ve loc i ty increment.
(2) Tota l ;"ai lable LH2 includes a 460 pound b ias .
The preceding t a b l e is a modification of t he res idua l propel lant assessment
provided i n the ASTP (SA-210) OT documentation, using the ac tua l FPR generated
i n t h i s a n a l y s i ~ .
SECTION 1
INTRODUCTION
Launch v e h i c l e performance i s p r e d i c t a b l e only b i t h i n c e r t a i n t o l e r a n c e s .
Therefore , dev ia t ions from a predic ted launch v e h i c l e t r a j e c t o r y a r e expected.
I n order t o e s t a b l i s h r e a l i s t i c dev ia t ion l i m i t s f o r t h e ASTP (SA-210) Launch
Vehicle Operat ional F l i g h t T r a j e c t o r y , a d i s p e r s i o n a n a l y s i s has been conducted
and is documented i n t h i s r e p o r t .
The nominal t r a j e c t o r y prescr ibed f o r t h i s a n a l y s i s is t h e ASTP (SA-210)
Launch Vehicle 500 Pound Launch Window Opening OT. This t r a j e c t o r y is docu-
mented i n Reference 1.
The e r r o r sources considered a r e those a s s o c i a t e d wi th p r e d i c t i o n s of
v e h i c l e c h a r a c t e r i s t i c s , v e h i c l e systems performances, and f l i g h t environment.
The nominal v e h i c l e , t h e boost t r a j e c t o r y s imula t ions , t h e e r r o r sources , t h e
a n a l y t i c procedures u t i l i z e d , and t h e r e s u l t s a r e d iscussed i n t h e fo l lowing
s e c t ions .
Launch v e h i c l e guidance system inaccurac ies were determined from t h e
guidance e r r o r a n a l y s i s , which is documented i n Volume I1 of t h i s pub l i ca t ion
(Reference 2) . These d a t a a r e composed of i n d i v i d u a l e r r o r source t r a j e c t o r y
e f f e c t s , which a r e s t a t i s t i c a l l y combined t o provide t r a j e c t o r y parameter
d i spe r s ion envelopes. F ixed t ime s t a t e v a r i a b l e ca rds a r e provided wi th
Volume I1 t o f a c i l i t a t e o r b i t a l t r a j e c t o r y d i s p e r s i o n ana lyses .
SECTION 2
DISCUSS ION
2 .1 Mission Descr ip t ion
The Apollo Soyuz Tes t P r o j e c t (ASTP) is a j o i n t USA and USSR ven tu re
c o n s i s t i n g of s e p a r a t e Apollo and Soyue spacecra f t launches f o r an e a r t h
o r b i t rendezvous and docking. The Soyuz w i l l be launched f i r s t on J u l y 15,
1975 and i n s e r t e d i n t o a 1881228 km. (101.5/123.1 n.mi.) e a r t h o r b i t inc l ined
a t 51.78 degrees. Subsequently, t h e Soyuz o r b i t w i l l be c i r c u l a r i z e d a t
225 km. (121.5 n.mi.) . Approximately 75 hours a f t e r t h e Soyuz launch, t h e
Apollo s p a c e c r a f t w i l l be launched and i n s e r t e d i n t o a 1501167 krn. (81/90 n .mi . )
e a r t h o r b i t coplanar ~ 5 t h t h e Soyuz o r b i t . The Apollo w i l l then rendezvous
and dock wi th t h e Soyuz. The two s p a c e c r a f t w i l l remain docked f o r approxi-
mately two days, dur ing which t ime t h e crews w i l l exchange v i s i t s and opera-
t i o n a l procedures. A f t e r a d d i t i o n a l docking t e s t s , t h e s p a c e c r a f t w i l l
s e p a r a t e and conduct independent a c t i v i t i e s . The Soyuz w i l l deorb i t approxi-
mately 46 hours a f t e r t h e i n i t i a l undocking, and t h e Apollo w i l l remain i n
o r b i t , conducting experiments, f o r f i v e a d d i t i o n a l days.
2.2 Launch Vehicle and T r a j e c t o r y Descr ip t ion
The launch v e h i c l e and t y p i c a l t r a j e c t o r i e s a r e descr ibed i n Reference
1. Fea tu res p e r t i n e n t t o t h i s a n a l y s i s a r e d iscussed i n t h e fo l lowing sub-
s e c t i o n s . Associated d i s p e r s i o n da ta a r e d iscussed i n suSsect ions 2.3 and
2.4.
2.2.1 Launch Vehicle
The Apollo launch v e h i c l e is Sa tu rn I B 210. It i s composed of t h e
S-IE-10 f i r s t s t a g e , an i n t e r s t a g e , t h e S-IVB-210 second s t a g e , and t h e S-IU-210
Instrument Unit . Major spacecra f t elements a r e t h e 0 1 - 111 Col~~~.~and Module, t h e
SM- 111 Serv ice Module, the SLA- 18 Spacecraf t Launch AdapLer anil t h e u>!-2
Dockin& bioaule. A Launch Escape Systelkl (LES) completes t h e space v e h i c l e . A
v e h i c i e weight breakdown is presented i n Table 1.
2.2.2 F ! igh t Environment - The 1963 P a t r i c k A i r Force Base atmosphere model, defined i n Reference
3, i s t h e nominal atmosphere used i n t h i s a n a l y s i s . The nominal wind i s t h e
J u l y mean p r o f i l e from Reference 4 supplemented by compatible data from
Reference 5 f o r a l t i t u d e s g r e a t e r than 27 k i lomete r s .
2.2.3 F l i g h t Sequence of Events
The nominal f l i g h t sequence of e v e n t s , f o r t h i s a n a l y s i s , is presented
i n Table 2 . O f f nominal propuls ion systems performances produce s i g n i f i c a n t
sequence changes. Of primary i n t e r e s t a r e t h e events which e s t a b l i s h Launch
Vehicle D i g i t a l Computer (LVDC) time bases and t h u s t h e subsequent events
dependert on t h e s e time bases. A discuss ion of p e r t i n e n t time bases and
a s s o c i a t e d even t s fo l lows.
1 ) Time Base 2 (T2) - Establ ished by S-IB s t a g e p rope l l an t l e v e l
sensor a c t u a t i o n i f a downrange v e l o c i t y 1 500 mlsec e x i s t s .
S i g n i f i c a n t dependent events a r e Inboard Engine Cut-Of f S i g n a l
(IECO), in te rconnec t ion of t h r u s t O.K. swi tches and f u e l dep le t ion
probe arming.
2 ) Time Base 3 (T3) - Establ ished when a n Outboard Engine Cut-Off
S igna l (OECO) is received by t h e LVDC due t o e i t h e r UIX o r fue l
d e p l e t i o n ; o r , by a backup LVDC s i g n a l i n i t i a t e d 13.00 seconds
a f t e r es tabl ishment of Time Base 2 . P e r t i n e n t dependent events
a r e u l l a g e rocket f i r i n g , S-IB re t ro - rocke t f i r i n g , S-IBIS-IVB
separa t ion s i g n a l , 5-2 engine s t a r t s i g n a l , I G M guidance i n i t i a t i o n ,
and Engine Mixture Ra t io (DIK) changes.
3) Time Base 4 (T4) - I n i t i a t e d approximately 0 . 2 seconds a f t e r
Guidance Cutoff S igna l (GCS). I n t h i s a n a l y s i s , GCS is r e -
ceived when t h e S-IVB s t a g e ob ta ins t h e t a r g e t v e l o c i t y l e s s t h e
p red ic ted v e l o c i t y increment from 5-2 t h r u s t decay. The s i g n i f i -
cant events subsequent t o T4 a r e t h e preplanned o r b i t a 1 maneuvers
and S-IVB s t a g e ven t ings .
It should be noted t h a t T2 and T3 es tab l i shments a r e nominally de-
pendent upon p r o p e l l a n t l e v e l sensor a c t u a t i o n s and p rope l l an t dep le t ion
de tec t ion . There fo re , es tabl ishments of these time bases a r e ve ry s e n s i t i v e
t o propuls ion system p e r t u r b a t i o n s , which a f f e c t p rope l l an t f lowra te , and
t h u s , tank l e v e l h i s t o r i e s ,
2 . 3 Dispersion Er ro r Sources
Vehic le manufacturing t o l e r a n c e s , p red ic ted system performance in -
a c c u r a c i e s , f l i g h t environment anomalies, and guidance hardware inaccurac ies
a r e sources of e r r o r s which s i g n i f i c a n t l y a f f e c t t r a j e c t o r y p r e d i c t i o n s . To
f a c i l i t a t e s t a t i s t i c a l ana lyses of such e r r o r e f f e c t s , t h r e e sigma t o l e r a n c e s
have been e s t a b l i s h e d . The t h r e e sigma t o l e r a n c e s considered i n t h i s a n a l y s i s ,
wi th corresponding r e f e r e n c e s , a r e d isplayed i n Table 3.
The LOX and RP-1 d e n s i t y cases presented he re in were g e ~ ~ e r a t e d from
J u l y propulsion p red ic t ions u t i l i z i n g t h e t apes de l inea ted i n Reference 11.
The t h r e e sig11:a wind data u t i l i z e d a r e t h e Reference 5 annual wind p r o f i l e s .
2.4 T r a j e c t o r y Dispersions and Ana ly t i ca l Procedures
T l~e t r a j e c t o r y parameter p e r t u r b a t i o n s r t ~ s u l t i n g frolr t h i s a n a l y s i s
a r e a ssu~ i~ed t o be rantlom, independent, and normally d i s t r i b u t e d . These
assu~npt ions al low a p p l i c a t i o n of t h e Root-Sum-Square (RSS) s t a t i s t i c a l con2-
b i n a t i o n method t o produce a reasonable t r a ; ec to ry d i spe r s ion envelope.
Dispersed t r a j e c t o r i e s were generated wi th each of t h e t t l ree sigma
to le rances de l inea ted i n Table 3. E f f e c t s on p e r t i n e n t t r a j e c t o r y parameters
a t S-IB/S-IVB s t a g e s e p a r a t i o n and o r b i t i n s e r t i o n were determined and combined
a s fo l lows:
+ RSS = JZ(+' ;
- RSS ,/- ; where
AP = perturbed parameter - nominal parameter.
These RSS values d e f i n e a reasonable t h r e e sigma f l i g h t envelope f o r
t h e ASTP (SA-210) Launch Vehic le Operat ional F l i g h t T r a j e c t o r y . In a s i m i l a r
manner, u t i l i z i n g t r a j e c t o r y d i spe r s ion d a t a , t h e S-IVB F l i g h t Performance
Reserve (FPR), r equ i red t o o f f s e t t h e combined t h r e e sigma d e v i a t i o n s , was
determined, Th i s FPR and o t h e r t r a j e c t o r y d i s p e r s i o n r e s u l t s a r e presented
i n Sect ion 3.
SECTION 3
RESULTS
3.1 T r a j e c t o r y Dispers ions
T r a j e c t o r y d i s p e r s i o n data a r e presented f o r two e v e n t s , S-IBIS-IVB
s t a g e s e p a r a t on and o r b i t i n s e r t i o n . Table 4 p resen t s t h r e e sigma t r a -
j e c t o r y parameter dev ia t ions produced a t S-IBIS-IVB s e p a r a t i o n by t h e S-IB
s t a g e propuls ion, non-propulsion, and f l i g h t environment pe r tu rba t ions .
Table 5 provides s i m i l a r da ta derived from S-IVB s t a g e p e r t u r b a t i o n s . 'Fables
7 and 8 d i sp lay corresponding data a t o r b i t i n s e r t i o n . I n t h e cvent t h a t
both - + t h r e e sigma p e r t u r b a t i o n s of t h e same e r r o r source produce e f f e c t s
wi th l i k e a l g e b r a i c s i g n , only t h e l a r g e r e f f e c t i s included i n t h e RSS.
Tables 6 and 9 d i s p l a y p red ic ted t h r e e sigma f l i g h t envelopes a t
S-IBIS-IVB s e p a r a t i o n and a t o r b i t i n s e r t i o n , r e s p e c t i v e l y . These envelopes
a r e t h e root-sum-square of t h e p rev ious ly mentioned e r r o r source group e f f e c t s
with t h e RSS of t h e I n e r t i a l Measurement Unit (IMU) e r r o r e f f e c t s included i n
Table 9 . Ind iv idua l IMU e r r o r e f f e c t s a r e provided i n Reference 2 .
Resu l t s of t h e a n a l y s i s show t h a t t h e expected extreme dev ia t ions f o r
T2 a r e +2.70 and -2.53 seconds. Analysis a l s o r e v e a l s t h a t t h e maximum
dev ia t ions expected f o r T 3 a r e +2.82 seconds and -2.65 seconds. Since S-IB/
S-IVB s t a g e s e p a r a t i o n , 5-2 i g n i t i o n , and I G M i n i t i a t i o n t imes a r e dependent
on T3, the maximum expected dev ia t ions f o r t h e s e events a r e t h e same a s those
of T3. Th i s f a c t is r e f l e c t e d i n Table 6 f o r S-IBIS-IVB s t a g e s e p a r a t i o n .
A b a s i c c r i t e r i o n f o r Launch Escape System (LES) j e t t i s o n i n g is t h a t
dynamic p ressure (q) haa decreased t o one pound pe r square foo t ( p s f ) .
Therefore , t h e t h r e e sigma d i spe rs ion on t h e time t h i s occurs was determined
t o f a c i l i t a t e s e l e c t i o n of a s a t i s f a c t o r y LES j e t t i s o n t ime. It was found
t h a t q = 1 psf may occur a s e a r l y a s T 3 +20.81 seconds or a s l a t e a s T3
+31.01 seconds. These extremes r e f l e c t dev ia t ions of -4.83 seconds, and
+5.37 seconds, r e s p e c t i v e l y , from t h e nominal T3 + 25.64 seconds. Thus,
cu r ren t OT s imulat ion LES j e t t i s o n t ime of T3 + 32 seconds provides 3 u
p r o b a b i l i t y t h a t q 5 1 p s f .
The S-IVB s t a g e ENR s t e p down is a T3 even t , t h e r e f o r e the expected
dev ia t ion extremes a r e those presented previously f o r T3. It i s found t h a t
t h e maximum expected v a r i a t i o n s i n T4 a r e f10.99 seconds and -10.46 seconds
a s shown i n Table 9 . These v a r i a t i o n s a r e p r imar i ly due t o S-IVB propuls ion
pe r tu rba t ions .
The e r r o r sources prescr ibed f o r t h i s a n a l y s i s , Table 3 , do not include
cond i t ions and t o l e r a n c e s which c o n t r i b u t e t o a r e a l i s t i c v e h i c l e a t t i t u d e
r a t e envelope determinat ion a t S-IBIS-IVE phys ica l s e p a r a t i o n o r o r b i t i n s e r t i o n .
Consequently, t h e t o t a l a t t i t u d e r a t e envelopes have been omitted f r m Tables
6 and 9 .
Durin, S-IVB s t a g e f l i g h t , r o l l c o n t r o l i s maintained by t h e Auxi l i a ry
Propulsion System (APS). Th i s system a l s o assumes p i t c h and yaw c o n t r o l a t
T4 + 3.5 seconds, E s s e n t i a l l y , t h e APS c o r r e c t s a t t i t u d e e r r o r s when t h e
a t t i t u d e e r r o r s i g n a l s exceed one degree. These c r i t e r i a a l low a t t i t u d e e r r o r s
t o approach 2 one degree a t o r b i t i n s e r t i o n , thus a two degree APS deadband
e x i s t s . Consequently, a t t i t u d e d i f f e r e n c e s (d i spe r sed - nominal) could be
increased by n e a r l y two degrees due t o t h e APS deadband i f t h e dispibrsion and
t h e nominal a t t i t u d e e r r o r s approached oppos i t e deadband l i m i t s . Such a,>
inc rease would be compounded by t h e RSS process , t : ,e reforc , tl-,is met' - ' is
not a p p l i c a b l e f o r a t t i t u d e d i spe r s ion envelope d e r i v a t i o n s a t or.:>i: i6.1.
Accordingly, a t t i t u d e envelopes have been excluded from Table 9 . These
a t t i t u d e envelopes a s w e l l a s t h e a t t i t u d e r a t e envelopes d iscussed i n t h e
previous paragraph a r e c u r r e n t l y derived a t MSFC by a n a1 t e r n a t e method.
Three sigma d i spe rs ion envelopes of p e r t i n e n t design parameters dur ing
S-IB s t a g e f l i g h t a r e displayed i n Table 10. Tables 11 and 12 provide p e r t i -
nent performance t r ade-of f f a c t o r s a t S-IBIS-IVB s e p a r a t i o n and o r b i t i n s e r t i o n ,
r e s p e c t i v e l y . Table 13 e x h i b i t s t h e e f f e c t s a t o r b i t i n s e r t i o n of l a r g e
guidance pla t form azimuth misalignments. Such misalignments may r e s u l t kom
ground c o n t r o l equipment inaccurac ies i n t h e event t h a t a backup alignment
scheme is employed. These e f f e c t s a r e not included i n t h e t h r e e sigma enve-
lopes presented h e r e i n .
3.2 S-IVB Stage F l i g h t Performance Reserve
The S-TVB s t a g e 3 u F l i g h t Performance Reserve (FPR) requirements f o r
t h i s mission a r e derived i n Table 14. The requirements a r e 1172 pounds of
LOX and 683 pounds of LH2. A previous a n a l y s i s , documented i n Reference 19,
has e s t a b l i s h e d t h a t FPR v a r i a t i o n wi th in a 700 pound launch window i s l e s s
than 50 pounds. Therefore , t h i s FPR is considered v a l i d a t any point i n t h e
s p e c i f i e d 500 pound launch window. U t i l i z i n g t h i s FPR, t h e r e s i d u a l
S-IVB p rope l l an t assessment presentefi i n Reference i i s inodified i n tlie
fo l lowing t a b l e .
LOX LH2 T o t a l (Pounds) (Pounds) (Poul~ds)
T o t a l on board a t GCS
(1 ) Unuseable
(2) T o t a l Avai lable
3 sigma FPR a l l o c a t i o n 1172 683 1855
Retna in ing launch window a 1 l o c a t ion 334 -,
7 0 - 404 ( 4 . 8 : l EMR) -
T o t a l a l l o c a t i o n 1506 753 2259
Excess a v a i l a b l e over a l locat io . . 262 9 1 3 53
Excess useable a t 4 . 8 : l EMR
Excess b i a s
(1) Unuseable determined by MSFCIMDAC t o a s s u r e t h e requ i red 6 . 7 m/sec dep le t ion cutoff t h r u s t decay v e l o c i t y increment.
( 2 ) T o t a l a v a i l a b l e LH2 include, a 463 pound b i a s .
Table 14 a l s o concains s i g n i f i c a n t i n d i v i d u a l p e r t u r b a t i o n e f f e c t s
on t h e S-IVB s t a g e p r o p e l l a n t components consumed and t h e r e a d i l y a p p l i c a b l e
t r ade -of f f c c t o r s . U t i l i z i n g t h e proper a l g e b r a i c s i g n , t h e s e t rade-off
f a c t o r s provide quick e s t i m ~ t e s of p e r t u r b e t ' o n e f f e c t s on S-IVB p r o p e l l a n t s
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1. CCSI) TN-FT- 74-35, ASTP (SA-210) Launch V e h i c l e O p e r a t i o n a 1 F l i g h t Tra- j e c t o r y , P a r t IIT, F i n a l Documenta t io~ . , d a t t ~ d .Ta~! t~ary 2 1 , 1 4 7 5 : a s u p d a t e d by CCSD L e t t e r , R . M . B l a c k s t o c k t ,) J . L. C r a f t s . d a t e d 1 . r ; . r-uary 7 , 1975.
2 . CCSD TN-FT-75-43, :lSTP (SA-210) Lnuncti Vehic l v g ) ~ ? r a t i o n a l Flirrh: T r a i e c t o r k , D i s n e r s i o n A n a l v s i s . Volc .e :!. G ~ ~ i d a n c t Ha idwar t E r r o A n a l v s i s ( 1 ) d a t e d A p r i l 4 . 1 9 i j . i *
3. NASL\/>ISFC TPR-53139, ti R e f e r e n c e ~ \ t n o s p h e r e f o r P a t r i c k :ll:Y, F l o r i d a , - Annual (1903 R e v i s i o n ) , d a t e d S e p t t ~ ~ ~ b e r 23 , 1964 .
4 . NASA/MSFC S&E-rlEKO-YT-77-71, S u b j e c t : Elon t t~ ly V e c t o r l lean Winds V e r s u s A l t i t u d e f o r Capr Kennedy, F l o r i t l a , f o r S k y l a b (INT-21) Wind B i a s 'I'rn i e c t o r u A n a l \ s i s , d a t e d J a n u a r y 1 8 , 1971.
5 . NASA/MSFC Rf -53956, Cape Kennedy Wind Component S t a t i s t i c s ? i o n t h l y and. Annual R e f e r e n c e P e r i o d s f o r A l l F l i g h t Azimuths from 0 t o 70 Kk! A l t i t u d e , d a t e d O c t o b e r 9 , 1969.
6 . NASA/MSFC R-AERO-F-27-67, S u b j e c t : S-IB, S-IVB T h r e e Sigma T o l e r a n c e Enve lope f o r Use i n t h e S t a g e I n c e n t i v e P l a n f o r V e h i c l e s AS-207 t h r o u g h AS-212, d a t e d F e b r u a r y 1 0 , 1967 (11).
7 . CCSD TN-A?-68-31', SA-?06/UI and SA-207/CSM ?.erodynamic A x i a l F o r c e C h a r a c t e r i s t i c s - F l i s s i o n 276, d a t e d March 1 , 1 9 6 t
8. NASA/MSFC R-P&VE-VAW-6b-119, S u b j e c t : S a t u r n I B T!~ree-Sipma R a d i a l C e n t e r of G r a v i t } D e v i a t i o n D u r i n g f i r s t S t a g e F l i g h t , d a t e d November 30, 1966 .
9 . NASA/MSFC S&E-AEXU-Y'r-91-71, S u b j e c t : Computer S u b r o u t i n e s f o r t h e Cape Kennedy Hot and Cold Atmospheres , 1971, d a t e d J u l y 2 3 , 1971.
l o . NASA/MSFC S&E-ASTN-SAB ( 7 1 - 9 ) , S u b j e c t : S-IR S t a g e P r o p u l s i o n Sys tem D i s p e r s i o n s f o r S k y l a b M i s s i o n s , d a t e d May b , 1971 .
11. CCSD TR-P&VE-75-222, F i n a l Launch V e h i c l e P r o p u l s i o n Sys tems F l i g h t Per fo rmance P r e d i c t i o n f o r SA-210, d a t e d J a n u a r y 3 1 , 1975.
12. NASA/MSFC Computer Card Decks 513A (Rev. 2 ) , 513B (Rev. l) , 51% (Rev. 1) t h r o u g h 515G (Rev. l ) , and 515H t h r o u g h 515s .
13 . NrSSr\,'MSFC R-P&VE-PPE-bv-M-99, S u b j e c t : S-IB S t a g e 200K and 205K H-1 Engine T h r u s t DecL1y P r o f i l e s , d a t e d J u n e 3, 1966 .
REFERENCES (CONTINUED)
NASA/MSFC SU-AERO-MFG-138-70, S u b j e c t : S ign Convention t o be Used i n D i s p e r s i a n Ana lys i s of ST-124M P la t fo rm Hardware E r r o r s f o r S a t u r n V and S a t u r n I B V e h i c l e s , d a t e d November 1 0 , 1970.
NASA/MSFC S&E-ASTR-SG-36-69, S u b j e c t : ST-124N P la t fo rm Hardware E r r o r s t o be Used i n Per iorming a Hardware E r r o r Ana lys i s f o r t h e S a t u r n I B and S a t u r n V Launch V e h i c l e s , d a t e d September 25 , 1969 ( U ) .
NASA/MSFC S&E-ASTN-SAB (72-20), Sut j e c t : S a t u r n IS P e h i c l e Engine S t a r t and Shutdown Performance C h a r a c t e r i s t i c s P r e d i c t e d f o r Skylab and S u b s e q ~ e n t Mis s ions , da t ed December 12 , 19 7 2 .
MSFC/CCSD Telecon - R . Ba i l ey t o R B lacks tock , Sub jec t : ASTP (SA-210) L/V O p e r a t i o n a l T r a j e c t o r y Dispers i on A n a l y s i s , DRL 444-V4, December 17 , 1974.
NASA/MSFC GFD/Grouudrule Approva 1 S h e e t , Task : ASTP (SA-2 10) Sa r u r n I B Veh ic l e Operat iona - 1 F l i g h t T r a j e c t o r y D i spe r s ion Ana lys i s , DRL 444-V4, appr3.vc.,.l 12-13-74; n,:d ; i ev i s ion -1, approveu 2-3-75.
CCSD TN-AP-71-La2, S1:ylab/Saturn IF3 Launch Kindow Dispe r s ion A n a l y s i s , P a r t I. da t ed J u c e 30. 1971.
CCSD TN-FT- 74- 19, ASTP (SA-210) Launch Veh ic l e P r e l i m i n a r y O p e r a t i o n a l F l i g h t T r a j e c t o r y D i spe r s ion A n a l y s i s , Volume I, da t ed J u l v 15. 197L .
MSFC/CCSD Telecons - R. B a i l e y t o N . Wi l l iams , S u b j e c t : C ~ r r e c t i o n s t o ASTP (SA-210) L/V OT Dispe r s ion Ana lys i s GFD, December 16 , 1974 and December 18 , 1974.
TABLE 1
ASTP (SA-210) L/V OPERATIONAL FLIGHI' TRAJECTORY 1)ISPERSION ANALYSIS VEHICLE WEIGHT BREAKDOWN
500 POUND LAUNCH WINDOW OPENING TRAJECTORY (POUNDS)
DM SM CM SLA Papels SLA (Fixed) Instrument Unit
kS-IVB Stage * W e a b l e S-IVB Prope l l an t
Orbi t I n s e r t ion Weight
I,OX Vented .. -2 Thrust Decay and Drain Prope l l an t S-IVB Cutoff Weight
S-IVB Prope l l an t Consumed S-IVB APS Prope l l an t Consumed LES Ullage Cases
S-IVB "90% Thrust" Weight
S-IirB GH2 S t a r t Tank S-IVB Buildup Prope l l an t Consumed Ullage Prope l l an t Consumed
S-IVB Weight a t Separa t ion
S-IVB Aft Frame Hardware E-IB/S .IVB I n t e r s t a g e S-12 Dry Weight S-IB Residuals and Reserves S-IVB Detonation Package S-IVB F r o s t Diss ipated S-IB F r o s t Diss ipated S-IB Sea l Purge Consumed (N2) S-IB Fuel Addi t ive Consumed (Oronice) S-IB Gearbox Co1;sumpt ion (RP- 1) Inboard Engine Thrust Decay Prpt. Consumed Outboard Engine Thrust Decay Prpt. Consumed
t o Separat ion S-IB Mainstage Prope l l an t Coneumed
Vehicle Weight a t F i r s t Motion
* Includes s u f f i c i e n t useable p r o p e l l a n t t o a s s u r e t h e r e q u i r e d 6.7 m / s dep le t ion t h r u s t decay v e l o c i t y increment-and a 460 pound LH2 b i a s .
* Composed of 1 ,434 pounds of LOX and 299 pounds of LH2.
TABLE 2
ASTP (SA- 2 10) LAUNCH VEHICLE OPERATIONAL FLIGHT TRAJECTORY FLIGHT SEQUENCE OF EVENTS
500 P O W LAUNCH WINDOW OPENING TRAJECTORY
LVDC FLIGHT TIIG FLIGHT
PROGRAM (HR: M I N : SEC) (SEC) TIME(SEC)
Guidance ke fe rence Re lease (GRR) ; I n i t i a t i o n of T i r e Base 0 . Title f o r S-75 T83iri@tage Igni* ; - - . Hold Down A r m Release S i g n a l . F i r s t Motion. Li f t -Of f S i g n a l ; l n ' t i a t e Time - Base 1 . I n i t i a t e P i t c l i and Ro.1 Maneuvers. Mach One. Maxi~~um Dynamic P re s su re . Con t ro l Gain S w i t ~ h P o i n t . Con t ro l Gain Swirch P o i n t . Enable S-IB P r o p e l l a n t Level Senso r s . A r r e s t A t t i t u d e Cccmands. Level Sensor Ac tua t ion ; I n i t i a t e Time Base 2 . Inboard Engine Cutoff (IECO). Outboard Cngi;-.e Cutoff (OECO); I n i t i a t e T ine Base 3 . U l l age Rockets I g n i t i o n . s e p a r a t i o n s i g n a l . S-IB/S-IVE P h y s i c a l .5 , p a r a t i o n . 5-2 Engine S t a r t Command. 90X 5-2 T h r u s t Level . Command i . > : l F I R . Ul lage Burn Out . J e t t i s o n Ul lage Rocket ? lo tors . Dynamic P r e s s u r e = 1 PSF. LES J e t t i s o n . Command Ac t ive Guidance I n i t i a t i o n . Con t ro l G a i n Switch P o i n t . Con t ro l Gain Switch P o i n t . Comniand DlR S h i f t t o 4 . 8 : l . Guidance Cutof f S i g n a l (GCS). I n i t i a t e Time Base 4 ; I n e r t i a l A L ~ i t u d e F r e e z e . I n i t i a t e LO!' :VV. O r b i t I n s e r t i o n . I n i t i a t e LH2 hTV. I n i t i a t e a maneuver t o a l i g n and ma in t a in t h e S-IVB/CSM a l o n g t h e l o c a l h o r i z o n t a l , nose l e a d i n g , p o s i t i o n 1 down. End LOX WIT. End of t r a j e c t o r y s i m u l a t i o n .
GROUP -
TABLE 3
ASTP (SA-210) L/V OPERATIONAL FLIGHT TRAJECTORY DISPERS'ION ANALYSIS THREE SIGMA TOLERANCES
ITEM - S-IB Stage Non-Propellant Mass Non-Propulsion Thrust Misalignment (Pi tch)
Thrust Misalignment (Yew) Thrust Misalignment (Roll) Axial Force Coeff ic ient Axia 1 Force Coefficient
*Center of Gravity Offset ( y ) *Center of Gravity Offset (2)
Environment
S-IB Stage Propulsion
REF i RENC E
Headwind Tailwind Right Cross Wind Left Cross Wind Atmosphere Atmosphere
High LOX Density Low LOX Density High Fuel Density Low Fuel Density Fue 1 llass Fuel Mass LOX Mass W?X Mass Thrust and Flowrate IS? and Flowrate Engine Mixture Ratio Engine Mixture Ratio
+ 310 Pounds - + 0.62 Degrees - + 0.62 Degrees - + 0.62 Degrees - Maxi mum Minimum + 0.05 Meters - + 0.05 Meters -
I Annua 1 3 cr where 3
Annua 1 ava i lab le , 3
Annual maximum 3
Annua 1 otherwise 3
Hot Atmosphere P r o f i l e 9 Cold Atmosphere P r o f i l e 9
- 3 u Ju ly Surface Winds + 3 u Ju ly Surface Winds - 3 u Ju ly Surface Temp. + 3 u Ju ly Surface Temp. + 0.60% - 0.60% + 0.45% - 0.60% + 1.5 '/. - + 1.95 Seconds - + 2800 Pound Max. Residual - 1550 Pound Min. Residual
H- 1 Engine Thrust Decay - + RSS of 22.5% of Nominal 13 Thrust Decay Impulse For Each H-1 Engine
S-IVB Stage Non-Propellant Mass - + 200 Pounds Non-Propulsion *Center of Gravity Offset (y) - + 0.05 Meters
+ 0.05 Meters *Center of Gravity Of £set (e) - Thrust Misalignment (Pi tch) - + 1.24 Degrees Thrust Misalignment (Yaw) - + 1.24 Degrees
Instrument Unit I n e r t i a l Measurement Units --- 14 & 15
* Referenced t o Project Apollo Standard Coordinate System 9.
TABLE 3 (CONTINUED)
ASTP (SA-210) L/V OPERATIONAL FLIGHT TRAJECTORY DISPERSION .4NALYSIS THREE SIGMA TOLERANCES
ITEM - TOLERANCE REFERENCE
S-NB Stage 5-2 Thrust Decay Dispersion Limits 16 & 18 Propulsion Cases 1 through 12 I d e n t i f i e d below
S-IVB Stage Engine Performance Dispersions
* Nominal ASTP (SA-210) Launch Vehicle Operat ional F l i g h t Tra jec to ry Propulsion Data.
MSF C Tape No.
00029
00169
00254
00562
00701
00758
00916
0 10 59
*28311
*28311
3.28311
*28311
Case
1
2
3
4
5
6
7
S
9
10
11
12
Deviations from Nomina 1
LH2 Flowrate ( l b s l s e c )
+O .830
-0.830
M.396
-0.396
+ l . 083
-1.083
M.253
-0.253
M . 6 4 5% LOX Load
-0.64 57; LOX Loa d
+0.902% Fuel Lead
-0.9022 Fuel Load
LOX F lowra t e ( l b s l s e c )
+7.526
-7.526
-14.659
-0.659
-0.167
t0.167
+3.053
-3.053
Tota 1 Flowrate ( l b s l s e c )
+8.356
-8.356
+ l . 055
- 1.055
10.917
-0.917
+3.306
-3.3C6
Engine Mixtdre
Ra t io
H . 0 3 5
-0.035
-0.018
M.018
-0.072
M . 072
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-0.020
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( s e c s )
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-0.233
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+3678.
-3678
+1.935 1 +1504.
-1.935 -1504.
-0.074
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N . 0 9 7
+ 349.
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