Launch Vehicle No. 6 Flight Evaluationx

8/8/2019 Launch Vehicle No. 6 Flight Evaluationx

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N A S A - M S C - G - R - 6 5 - 5an d M S C - G - R - 6 6 - 2S u p p l e m e n t a l R e p o r t 23 3 3

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.

P R E P A R E D B Y

X67- U 13 5(ACCESSION NUMBER)

LAU NCHVEHICLE NO. 6

FLIGHTEVALU AT I ON (U )

U . S . G o v e r n m e n t A g e n c i e s On l y

Engineering Rep or t 13227- 6

(THRU)

fiL ' . .(NASA CR ORT-MXOR AD NUMbrftRJ( N A S A - C R - 8 3 0 9 0 ) L A U N C HE V A L U A T I O N ( M a r t i nCo.)

F e b r u a r y1966N O . 6 F L I G H T N 7 5 - 7 5 6 2 3

Unclas00/98 23571


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C o p y No.


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ER 13227 -6 N A S A - M S C - G - R - 6 5 - 5an d MSC-G-R-66-2Supplemental Report 2

F e b r u a r y 1966

Published as Supplemental Report 2to: G e m i n i Program Mission ReportsG e m i n i V I a n d Gemini VI- AMSC-G-R-65-5 an dMSC-G-R-66-2by: G e m i n i VI and VI- A Mission Evaluation TeamNational Aeronautics and Space AdministrationM a n n e d Spacecraft CenterH o u s t o n , Texas

LAU NCH VEHICLE NO. 6FLIGHT

EVALU AT I ON (U )

Approved by

oUL. J. Rose

Assi st ant Technical DirectorT e s t Evaluation

CurlanderDirector

IN ANr WM'ilEftPROHiBlTtO B Y

.V ,V IIS'C C O f F N T SP E R S O N i3

Prepared byMARTIN COMPANY, BALTIMORE DIVISION

Baltimore, Maryland 21203Under CONTRACT AF 04(695)-394

PRIORITY DX-A2

ForSPACE SYSTEMS DIVISION

AIR FORCE SYSTEMS COMMANDUNITED STATES AIR FORCE

Lot Angele, California


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ii

FOREWORD

This report has been prepared by the Gemini Launch Vehicle Pro-gram Test Evaluation Section of the Martin Company, Baltimore Divi-sion. It is submitted to the Space Systems Division, Air Force SystemsCommand, in compliance with Contract AF04(695)-394.

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I ll

CONTENTSPage

Foreword nSummary . . . vii

I Introduction . .. I ~lII. System Performance - 1

A. Trajectory Analysis . . . . II- 1B. Payload Capability 11- 40C. Staging - 40D . Weight Statement 11-41

III. Propulsion System - 1A. Launch Attempt (12 December 1965) - 1B. Engine Subsystem I I I - 2C. Propellant Subsystem - 24D . Pressunzation Su bsystem - 70E. Environm ental Contro l Ill- 81

IV. Flight Control System IV- 1A. Stage I Flight IV- 1B. Stage II Flight . . IV- 9C. Post- SECO Flight IV- 9

V. Hydrau lic System . V- lA. Stage I . . . . V- lB. Stage V- 5

VI. Guidance Systems VI - 1

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IV

C O N T E N T S ( c o n t i n u e d )Page

A . R a d i o G u i d a n c e S y s te m P e r f o r m a n c e V I - 1B . S p a c e c r a f t Inertial G u i d a n c e S y s te m A s c e n tP e r f o r m a n c e V I - 6

VII. Electrical S ystem A na lys i s VII -1A . C o n f i g u ra t i o n V I I - 1B. C o u n t d o w n an d Flight P e r f o r m a n c e V I I - 1

VIII. I n s t ru m e n t a t i o n S y ste m V I I I- 1A . A i rb o rn e I n st ru m e n t a t i o n V I I I- 1B . La n d l m e I n s t ru m e n t a t i o n V I I I- 3

IX . R a n g e S a f e ty and O r d n a n c e I X - 1A . C o m m a n d C o n tr o l R e c e i v e r s IX - 1B. M I S T R A M I X - 1C . O r d na n c e I X - 2

X . Ma lfunc t io n Detec t ion S ys tem X - lA . C o n f i g u r a t i o n X - lB . S y st em Pe r f o rm a n c e X - 2

X I. C r e w S a fe ty X I - 1A . G T - 6 A L a u n c h A t te m p t X I - 1B . P r e l a u n c h W i n d s O p e r at io n s X I - 1C . Slow M al fu n c t i on M o n i t o r i n g X I - 1 2

XII. A i r f ram e S ys tem XII-1A . S t ruc tu ra l Loads XII -1

E R 1 3 2 2 7 - 6


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C O N T E N T S ( c o n t i n u e d )

PageB. P O G O XII- 19

. A G E a n d Facilities 1 - 1A . Mechanical A G E XIII- 1B. Master Operations Control Set - 2C. Facilities XIII-3

XIV. Reliability XI V- 1XV. Range Data XV- 1

A. L a u n c h Attempt Data and Film Distribution . . . XV- 1B. Launch Data Distr ibut ion XV- 2C. Launch Film Coverage XV- 7

XVI. Prelaunch a n d C o u n t d o w n Operations XVI - 1A. Prelaunch XVI- 1B. Launch Att em pt C o u n t d o w n Summary XVI - 1C. Recycle and Prelaunch Activity X V I - 2D . C o u n t d o w n Summary X V I - 2

X V I I . Configurat ion Summary X V I I - 1A. Launch Vehicle Systems Description X V I I - 1B. Major C o m p o n e n t s X V I I - 3

XVin. References X V I I I - 1A p p e n d i x A : Summary o f G e m i n i Launches A - l

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Vl lRECEDING PAGE BLANK NOT FILMED.

S U M M A R YThe scheduled launch of G T- 6A on 12 December 1965 was terminatedd u e to a premature disconnection of tail plug 3D1M. Engine s h u t d o w n

was automatically initiated 1.24 seconds after MOCS T n by holdfireC- 4 (programmer reset monitor), w h e n the pro gramm er in the three-axis reference system (TARS), having been initiated by th e inad ver tentdisconnect of 3D1M, started prematurely. Engine s h u t d o w n was pre-ceded by a normal start transient in S/ A 1, whereas S/ A 2 exhibiteda n abnormal pressure buildup. Investigation of the abnormality led toa n inspection of the gas generator, which revealed that a plastic dustcover had been left in the S/ A 2 gas generat or assembly. All othersystems performed properly during the launch attempt.

O n 15 December 1965, G em ini- Titan 6A ( G T - 6 A ) was launched suc-cessfully and on schedule from Complex 19, Cape Kennedy, Florida.Launch vehicle / spa ce cr aft separation was completed 361 seconds afterl i f t o f f . Spacecraft r e - en t ry was accomplished after completion of 17. 1orbits.

The 240- m inut e c o u n t d o w n was picked up at 0529 EST on 15 Decem-ber and c o n t i n u e d without incident thr ough l i f t o f f at 0837 EST. Thespa cecr aft was inserte d into an elliptical orbit with a perigee of 87nautical miles and an apogee of 140. 4 nautical miles; all test objectivesfor the launch were achieved.Stages I and engines operated satisfactorily throu ghout p oweredflight. Stage I burning time was 160.4 seconds, with shutdown initiatedby oxidizer exhaustion. Stage II engine operation was terminated by aguidance com ma nd after 181.6 seconds of operation.The flight control system (FCS) m aintained satisfactory vehiclestability during Stages I and II flight . The primary FCS was in co m-mand throughout the flight. Vehicle ra tes during Stage I flight neverexceeded 1. 9 deg/ sec, and the maximum attitude error was 1. 7 degrees.The maximum rate and attitude error that occurred during staging didnot exceed 3. 7 deg/ sec and 2.8 degrees, respectively.P erformance of the radio guidance system (RGS) was satisfactory.Pitch and yaw steering signals and SECO discrete commands wereproperly executed.I G S pitch, yaw and roll performance for the entire flight appearednormal. The dispersions between IG S and pr imar y system attitudee r ro r s remained within acceptable limits during powered flight.

E R 13227- 6


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Vlll

The hydraulic system operated satisfactorily during launch opera-tions and both stages of flight. There were no significant pressureperturbations at l i f t o f f or duri ng flight.The electrical system f u n c t i o n e d as designed throughout the launchc o u n t d o w n and flight. Power transfer to vehicle batteries was smooth.All channels of the PCM instrumentation system f u n c t i o n e d satis-factorily throughout the fright. The landlme ins trumenta tion systemalso f u n c t i o n e d satisfactorily prior to and up to l i f t o f f . All airborneinstrumentation hold functions monitored in the blockhouse remainedw i t h i n specification throughout the c o u n t d o w n .The ordnance system umbilical drop weight release, explosive launchnuts and stage separation nuts operated as designed. The prevalveswere not replaced after the 12 December 1965 launch attempt, therefore,

the valves were open prior to propellant loading. The performanceso f the command control receivers and the M I S T R A M transponder weresatisfactory.Malfunction detection system (MDS) performance during preflightcheckout and flight was satisfactory. There were no switchover com-mands during the flight.The flight environment encountered by GT-6A was within design re-quirements. Flight loads were well w i t h i n the structural capabilitieso f the launch vehicle. The most critical loading ( w h i c h occurred at pre-BECO, aft of Station 320) reached 103. 6% of design limit load in com-

pression.The longitudinal oscillation (POGO) on GT-6A reached a maximumvalue at Station 280 of 0. 115 g zero-to-peak at frequencies of 13. 7 cpsat LO + 146. 8 seconds and 16. 8 cps at LO + 153. 9 seconds. This wasthe lowest POGO experienced on any Gemini f l i g h t to date.Crew safety monitoring, which was c o n d u c t e d at N A S A - M S C , wasactive during prelaunch and the launch. All guidance monitor parameterswere nominal and no corrective action was required during the flight.The precount operation progressed without problems. All AGE and

facilities operated without incident during the countdown. Propellantloading was completed within the scheduled time span and to the specifiedload and temperature limits.All electrical umbilicals disconnected in the planned sequence andwithin 0. 854 second. Engine blast and heat damage to the launch standwas minor.

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IX

G LV- 6 Test Objectives and ResultsObjective Results

PrimaryP - l

P - 2

P - 3

Demonstrate satisfactoryboost by the Geminilaunch vehicle system ofa manned Gemini space-cra ft into the prescribedorbital insertion condi-tions.

P - l Orbit insertion waswithin the predictedtolerances for V, hand .

Evaluat e lau nch vehiclesubsystem performanceduring powered flightfor mission success andcrew safety.

P - 2

Demonstrate the effec-tiveness of launch op-erations including prop-er operations of neces-sary ground/ range sup-port systems to achievethe prescribed rendez-vous mission launchrequirements.

-

All systems performedsatisfactorily through-out flight. The POGOoscillation (0. 115 gzero- to- peak) was thelowest encou ntered .G T - 6 A was erectedand ready for launchcountdown seven daysafter the launch ofG T - 7 from the samelaunch pad.

Secondarys - i

S- 2

Evaluate trajectory per-formance of the launchvehicle system for re-fining capability andpredictions fo r futuremissions.

s- i Vehicle flight waswithin the 3- sigm apredicted trajectory.

Demonstrate ability toload propellants toweight and t emperaturelimits imposed by pay-load and vehicle re-quirements.

S- 2 Tanks were loadedwithin the requiredtolerances of weightand t emperature .

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1-1

I. INTRODUCTIONThis report presents an engineering evaluation of Gemini LaunchVehicle No. 6 ( G L V ~ 6 ) systems performance during the 15 December1 9 6 5 c o u n t d o w n , launch and boost phase of the Gemini 6A mission.Discussion of pertinent aspects of the 12 December launch attempt hasalso been included.The GT-6A vehicle was launched from Complex 19, Cape K e n n e d y ,Florida at 0 8 3 7 hours EST on 15 December 1965, 12 days after the launcho f GT-7 from the same pad. A successful flight was achieved, and thespacecraft was inserted into an elliptical orbit with a perigee of 87nautical miles and an apogee of 140. 4 nautical miles. Gemini 6A was

the seventh mission and the f i f th manned flight in the Gemini program,with Astronauts Walter Schirra and Thomas Stafford aboard the Gemini6 spacecraft. The one-day mission, with its objective of rendezvousw i t h the Gemini 7 spacecraft, was completed successfully on 16 Decem-ber 1965.GLV-6 was delivered to Cape Kennedy on 2 August 1965 in prepara-tion for the Gemini 6 mission, a rendezvous and docking exercise withan Agena Target Vehicle (ATV), scheduled to begin on 25 October 1965.On that day, the GLV c o u n t d o w n and the mission were terminated shortlyafter the Atlas-Agena l i f t o f f w h e n the ATV failed to achieve orbit.Subsequently, the Gemini 6 mission plan was changed to that of arendezvous with the Gemini 7 spacecraft and redesignated as the Gemini6 A mission. For the redefined 6A mission, it was required to launchGT-6A from Complex 19 within eight to twelve days f o l l o w i n g the GT-7flight. Following the successful G T ~ 7 operation on 4 December 1965,the GT-6A vehicle was erected on Complex 19, and both vehicle andlaunch complex were readied for a 12 December flight. On this datethe c o u n t d o w n proceeded on schedule through T~0 and engine ignition,but an automatic shutdown occurred due to inadvertent release of a tailplug. The GT-6A vehicle was recycled to permit a 15 December launch,which was accomplished on time and without incident and which wasf o l l o w e d by the Gemini 6 rendezvous with the Gemini 7 spacecraft.Significant events and tests accomplished for GLV-6 at ETR appearin Table 1-1.

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X3VENTS

UGLV- 6 on dock, ETRErection of GLV- 6Subsystem revenfication (SSRT)Prespacecraft mate verification ( 1140M1cSpacecraft mechanical mateElectrical integrated interface va

CM^->

CM

CM

_HJoint guidance and control ( J G & C )Joint combined systems test (JCS

04CM_PJ1JPropellant tanking testFlight configuration mode test (F


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I I - 1

II. SYSTEM PERFORMANCEA. TRAJECTORY ANALYSIS

1. Orbit InsertionGemini Launch Vehicle No. 6 (GLV- 6) performed as predicted andinserted the Gemini 6 spacecraft into earth orbit well within the allow-able tolerance limits to permit rendezvous with the G T- 7 spacecraft .G LV- 6 was steered in the lateral plane during Stage I I flight to aset of ephemeris data referenced to the time of insertion (or targeting).The values of two of these targeting parameters = ~0. 18719584 x

10 r a d / se c and T~ = 53 , 741.21 875 seconds) are given here for record;there are no observed values of these parameters. The targeted and ob-served inclination angles were 28. 895 and 28. 97 degrees, respectively.The targeted wedge angle of 0.2002 degree was exceeded. The observedre sidual wedge angle was - 0.075 degree, which meant that the total wedgeangle steered was - 0.275 degree.

A comparison of the predicted and observed insertion conditions isgiven in Table II-1. In this table and in all succeeding references to apredicted (nominal) trajectory, the data wer e obtained from the G LV- 645- day pre launch rep or t (Ref. 10), updated to reflect the actual space-craft weight (7821 pounds), guidance constants, T- l hour wind and at-mospheric data, and the - 1.07% pitch and - 1 .4% roll programmerbiases. Th e observed trajecto ry para m eters are tho se der ived by theMartin Company from the Final GE Mod III-G 10 pps data. These datahave been smoothed and corrected for both refraction errors and sys-tematic biases by the General Electric Corporation before submittal tothe Martin Company.

TABLE II -1Comparison of Insertion Conditions a t SECO + 20 Seconds

Altitude (naut mi)Inert ial veloc ity (fps)Inertial flight pathangle (deg)

PredictedNominal87. 12825,7310.002

G E M o dIII-G87.26025, 728- 0.054

ObservedMinusPlanned+ 0. 132- 3- 0.056

PreliminaryTolerance0.346+30. 3iO.' 125

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inertial flight path angle, and geoc entric radius agreed very closelywith values derived from the Bermuda tracking data. BET data (whichare derived from M I S T R A M data) are presented in Table II- 6Table II - 6 pre sents a comparison of the GE Mod III-G 10 pps andthe 2 pps data. Starting with G T- 8, only the GE Mod - G 2 pps datawill be available.' Com pa ring the two sets of data on Table II- 6, it isconcluded that the GE Mod III-G 2 pps data are satisfactory.The actual, as well as the predicted, nominal trajectory is pre-sented in graphical form in Figs. I I -2 through 11-24. On these graphs,the nominal traject or y is that documented in Ref. 10, updated to re-flect the actua l spa cec raft weight (7821 pounds), guidance constants,T- l hour wind and atmospheric data, and the - 1.07% pitch and - 1. 4%roll programmer biases. Th e observed flight da ta wer e obtained fromthe Mod III-G 10 pps data, smoothed and corrected for refraction errors

and systematic biases.A list of the p rimar y tr acking sources with the tra ject or y tim e inter-val covered by each is contained in Table II- 7.

5. G eodetic and Weat he r ParametersSignificant geodetic and weather parameters are shown in Table II-8.The atmospheric pressure and temperature variation with altitude isdepicted in Fig. - 25. The pressure was essentially standard, whilethe temperature was slightly warmer than stand ard. Figure 11- 26 pr e-sents the altitude history of the magnitude and direction of the wind.

At low altitudes the winds were light, increasing to a peak of 86 knotsat 42, 500 feet. The wind was essentially a tail wind with a small com-ponent from the left of the trajectory.6. Look Angles

A n initial decoder pitch- down command of about 0. 10 deg/ sec ,lasting approximately 0. 5 second, was issued at LO + 168. 21 seconds.Following this, a 0. 56 deg/ sec pitch- down command was issued for ap-proximately 0. 3 second. Thereafter, the pitch steering command de-creased to appro xim ately 0.25 deg/ sec within 2 seconds.During this period, the maneuver resulted in decreasing angles ofattack as shown in Fig. - 20. The m axim um look angle in pitch (LAP )occurred at LO + 335 seconds, when it attained a value of 22.2 degrees.This maximum value was within the boundary existing at tha t time, asshown in Fig. 11-27. The corresponding look angle in yaw (LAY) wasalso within the established limitation (20 degrees), as shown in Fig.11-28. The m axim um value of LAY was 6. 3 degrees which occurred160 seconds after liftoff.

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-

TABLE II - 7Data Available for Trajectory Analysis

SourceAFETR

GE

NASA- MSC

TypeM I S T R A M posi-tion, velocityand accelerationFPQ-6 radarposition, veloc-ity and accelera-tionFPS-16 radarposition, veloc-ity and accelera-tionBETM o d III-G radarposition, veloc-i t ySpacecraft IGSaspect param-eters

StationValkaria IEleuthera II

MILA 19.18GBI 3.18Grand Turk 7. 18Patrick 0.183.16

(composite)Cape Kennedy

Flight Coverage(sec from r ange- 0)+ 65 to 383+ 65 to 383

+ 13 to +380+ 147.6 to + 379.1+ 329.1 to 491. 190 to 379+ 147.6 to +379.1

+ 6 6 to + 3 7 9LO to + 381

LO to +367.9

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II-11TABLE II-8

Geographic and Weather Conditions at LaunchLocation

SiteSite coordinates-

Latitude (deg)Longitude (deg)Pad orientation (deg)

Complex 19

28.507 N80.554 W84.903 true azimuth

WeatherAmbient pressure (psi)Ambient temperature (F)Dew point (F)Relative humidity (%)Surface w i n d -

Speed (fps)Direction (deg)

Winds aloft (max):Altitude (ft)Speed (fps)Direction (deg)

Cloud cover

14.71686797720 042,50014729 2 true azimuth0. 2 alt o-cumulus

Reference Coordinate SystemTypeOriginPositive X-axisPositive Y-axisPositive Z-axisReference ellipsoid

Martin reference coordinate systemCenter of launch ring, Complex 19Downrange along flight azimuthtangent to ellipsoidTo left of flight azimuth tangent toellipsoidand_LX-axisForms a right-handed orthogonalsystemFischer

LaunchInitial flight azimuth (deg)Roll program (deg)

81. 4 true azimuth3. 5 cw

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FIDENTIAL

I9!

BECO(157. 162 sec)"

- :?u

a- 4

1

1

Predicted nominal wind run 80-GT-6 (final)*G E Mod I I I - G final flight data*

^IncludesRawinsonde balloon dataCape Kennedy0739 EST, 15 December 1965

\J Predicted BECO(156.578 sec)

.

...

_.

I 10 .. .30 40 50 60 70 80 90 100 110Time from Liftoff (sec)

,120 130 140 150 160 170 180

ER 13227- 6

Fig. II-2. Inertial Velocity Versus T i m e : Stage I Flight


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P r e d i c t e d n o m i n a l w i n d ru n 8 Q - G T - 6 ( f ina l ) *G E M o d I I I - G f i n a l f l i g h t d a t a *

B E C O( 1 5 7 . 1 6 2 s ec)

I n c l u d e sR a w i n s o n d e ba l l oon d a t aCape K e n n e d y0 7 3 9 E S T , 15 D e c e m b e r 1 9 6 5 Pre d i c t e d B E CO( 1 5 6 . 57 8 s ec)

'

50 6 0 7 0 8 0 9 0 1 0 0 1 2 0 1 3 0T i m e f r o m Li f to f f ( s e c )

140 150 160 170 180

F i g . I I - 3 . Inertial Flight Pat h An g l e Ve r sus T ime : Stage I Flight

n n j 4^^!U[.li I I H L ALE R 1 3 2 2 7 - 6


27/253

240> - 14 - 'NFIDENTIAL

220

200

180

160

140- -

1 20:

100

80

!

Predicted nominal w i n d run 80-GT- 6 (final)*G E Mod I I I - G f i n a l flight data*

- 6 (final)* BECO( 1 5 7 . 162 sec)

i//./ /

130 140 150 160 170 180Time from L i f t o f f (sec)

^ fl?

Fig. II- k. Altitude (h) Versus T i m e : Stage I Flight

CONFIDENTIALE R 1 3 2 2 7 - 6


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400

Predicted nominal w i n d run 80-GT-6 ( f in a l ) *GE Mod III-G f in a l f l ig h t data*

"IncludesR a w i n s o n d e balloon dataCape K e n n e d y0 7 3 9 EST, 15 December 1965

'"

10 20

BECO( 1 57 . 162 sec)

Predicted BECO( 1 5 6 . 5 7 8 sec)i '

70 8 0 9 0Time from Lif to f f (sec)

160

II-15

Fig. II-5. Downrange Position Coordinate (XF) Versus Time: Stage I Flight

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- 16 CONFIDENTIAL5

'f

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Predicted BECO /(156.578 sec) '

j ; __ .20 30 10 50 60 70 80 90 100 120 130 140 150 160 " 170 180Time from L i f t o f f (sec)

Fig. II-6. Cross-Range Position Coordinate (Y ) Versus Time: Stage I Flight

fj 1* CONFIDENTIAL

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320

280

* 240

% 200- 160 -

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Pred i c ted nomir

*IncludesRawinsonde balloonCape Kennedy0739 EST, 15 Dece

120

BECO(157. 162 sec)

Pred icted BECO( 1 5 6 .5 7 8 sec)

Ti me fro m Liftoff (sec)100 110 120 130 140 150 160 170

Fig. II-T. Vertical Position Coordinate (ZF) Versus Time: Stage I Flight

_ 1 1A NFIDENTIAWE R 1 3 2 2 7 - 6


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11-18 " CONFIDENTIAL

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.

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70 8 0 9 0Tim e f ro m Lif to f f ( s e c )10 0 110 120 130 140 150 160 170

Fig . H-8 . Mach Number (M ) Versus Time: Stag e I Flight

CON FIDENER 1 3 2 2 7 - 6


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B E C O( 1 5 7 . 1 6 2 s ec)

:

P r e d i c t e d B E C O :( 1 5 6 . 5 7 8 s ec)

100 110 120 130 140 150 160

F i g . 11-10. A x ia l Fo r ce V er s u s T ime : Stage I Flight

W 1 D E N T I A LCONFIDENTIAL

ER 13227-6


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60

P r e d i c t e d n o m i n a l w i n d ru n 8 Q - G T - 6 ( f i n a l ) *GE M od I I I - G f in a l f l ig h t data*

- I n c l u d e sR a w i n s o n d e b a l lo o n da taC a p e K e n n e d y0 7 3 9 EST, 15 D e c e m b e r 1965

IDEN7W II-21-2-B EC O( 15 7 . 162 sec)"

P r e d i c t e d B EC O( 1 5 6 . 5 7 8 sec)

60 70 8 0 9 0Time f ro m Liftoff (sec)

100 110 120 130 140 150 16 0

Fig. II-11. Aerodynamic Heating Indicator Versus Time: Stage I Flight

HUM?!!ER 1 3 2 2 7 - 6


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1 1 - 2 2

II Pred i c ted n o m i n a l w i n d run 8 0 - G T - 6 ( f i n al ) *G E Mod I I I - G f i n a l f l i g h t data*

- ,-.

:.

20

10

- 10

- I n c l u d e sR a w i n s o n d e balloon dataC a p e K e n n e d y0 7 3 9 E S T , 15 December 1 9 6 5

BECO( 1 5 7 . 162 se c ) "

Predicted BECO( 1 5 6 . 578 se c )

-20

-3 0

-4 0 10 20 30 40 50 6 0 70 80 90 100Time from Liftoff (sec)

120 130 140 150 160 170

F i g . 11-12. Stage I Angle of Attack History

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W^^^^WP*^"23

Predicted nominal w i n d ru n 80-GT- 6 ( f i n a l ) *G E M od IH-G final f l i gh t data*

-

ato1

.-* -- -

-30

-40

.

-50

-60 30 50

''IncludesRawinsonde balloon dataCape Kennedy0739 EST, 15 Dece mber 1965

BECO( 1 5 7 . 162 sec)"

60 70 80 9C 100Time from Liftoff (sec)

110

Predicted BECO(156. 578 sec)

120 130 140 150 160 170 180

Fig. 11-13 Stage I Angle of Sideslip History

E R 1 3 2 2 7 - 6


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2

11-24 -

26

24

22

Pred i c ted nominal wind run 8 0-GT-6 (final)5G E Mod III-G final flight d ata*

^I n clu desRawinsonde balloon dataCape Kennedy0739 EST, 15 December 1965

SECO + 20 (358. 737 sec)

Pred i c ted SECO + 20 (356. 894 sec)

140 160 180 200 220 240 260 280 300 320 340T i m e f r o m Li f to f f (sec)

360 380 400 420 440 460

-_^k 'TiM Fig . II- lA. Resultant Inertia! Velocity (V ) V er s u s Time: Stage II FlightER 13227-6


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P r e d i c t e d n o m i n a l w i n d r u n 8 0 - G T - 6 ( f i n al ) *G E M od I I I - G f ina l f l igh t data*

..

R a w i n s o n d e b a l l o o n d a t aC a p e K e n n e d y0 7 3 9 EST, 1 5 D e c e m b e r 1965

S E C O + 2 0 (358. 737 sec)

Predicted S E C O + 2 0 ( 3 5 6 . 8 9 4 sec)

140 160 180 200 220 240 260 280 300 320Time f r o m L if to f f (sec)

340 360 380 400 420 440 460

Fig . II-15. Inertial Flight P a th A ngle ( ) V er s u s T ime : Stage II Flight

J G 8 N F I D E N T I A L ER 1 3 2 2 7 - 6


39/253

- 26

--

6 0 0 ,

56 0

5 2 0

48 0

44 0

40 0

2 360 !

320

280

; d n o m i n a l w in d r u n 8 0 - G T - 6 ( f i n a l )*I I I - G f i n a l f l i g h t data*

.

.

SE

"

SECO+ 20 (358.737 sec)

* I n c l u d e sR a w i n s o n d e b a l lo o n da taC a p e K e n n e d y0 7 3 9 EST, 1 5 D e c e m b e r 1965 Predicted SECO+ 20 (356. 894 sec)

140 160 180 200 220 240 260 280Time from L i f t o f f (sec)

30 0 320 340 3 6 0 3 8 0

Fig. II-16. Altitude Versus T i m e : Stage II Flight

(HIULNI HLER 13227- 6


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4.0,

_ 3CD

0

Q

ir ' S E C O + 20 (358 73 7 sec)

; . . _ . ; _ L-

P r e d i c t e d n o m i n a l w in d r u n 8 0 - G T - 6 ( f in a l ) *GEMIn"Gflnafllht daa*

_, '_ i J i L_ V ' ~ i P r e d i c t e d S E C O + 20 (356. 894 sec)[- - . - ..1 ' t^ ' . > ,. I n c l u d e s_'_L' ,'-_ _t R a w i n s o n d e b a l lo o n d a t a !' r ' "' ' ' C a p e K e n n e d y 0 7 3 9 EST, 15 D e c e m b e r 1965 \ '~

~ Jr - - , f - - . "Lb"T '" J _l ! j - i 1-- - I W- ;~ ! - V - - ! :i - ' - : ; ' t i - > . -' - : 1_,4_;__

.*Li : _ i-4lt.--L: i L'. "-_u-J " .J LJ_J_ ' i i..'. ' ._ L ' 'i ' ' '_ ! _ _ _ . . -> . ! . i " " ' ~ i

- - ; I z l t - -, . _ - - -240 260 280 300 320 340 360 38040 160 180 200 220 ' - u _ ' - j i .!;_^i'4 0 0 4 2 0 4 4 0

Time f r o m Li f to f f (sec)

Fig. 11-17- Dovnrange Position Coordinate (X_)V er s u s T ime: Stage II flight

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- 28 40 >

-:.

-

;.n--rt.I.- -'

-120

-140

-160l

Predicted nominal wind run80-GT- 6 (final)* |GE Mod III-G final flight data*

] ^IncludesRawinsonde balloon dataCape Kennedy0739 EST, 15 December 1965 h ;!

;"~1-liS EC O + 20 ( 3 5 8 . 7 3 7 sec)

-$$

r*:t:fc

i-. . . . . . , . { . - : [ '"-^u- ;- -

*, I\

**

... , -i i ; . . i . j :.!.:.ISf-

--T" - :; - ' ; - _ , i ' i" I

Predicted S E C O + 2 0 (356. 894 sec)

-200120 140 160 180 200 220 240 260 280 300 320 340 36 0 380 400Time f r o m Li f to f f (sec) 480

Fig. 11-18. Cross-Range Position Coordinate (Y )Versus Time: Stage Plight

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520

Pred i c ted no mi nal wind run 80-GT- 6 (final)*GE Mod III-G final flight data*

'

Rawinsonde balloon dataCape Kennedy0739 EST, 15 December 1965

SECO + 20 (358.737 sec)

Predicted SECO + 20 (356.894 sec)

4-,

140 160 180 200 220 240 260 280 300 320 340Time from L i f t o f f (sec)

38 0 400 420 440 46 0 480

Fig. 11-19. Vertical Position Coordinate (Zp) Versus Time: Stage Flight

* E R 1 3 2 2 7 - 6


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44/253


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- 32 - 650

50

aIX

-50

60 0

550nj

500

-100 450

-200

-250

- ~ - 'V

DQ -350-

-4001-

214. 7r

214.6

214.53>1 214.4

1.0

0.5

;II

5-kV

0.0

-0.5

214.3 -1.0

2 ,

: Cross -range velocity

\ Inertia! f l i gh t p a t h angleJLIiei L lc L i 1- *

! Cross range

i G r o u n d rang

SECO+ 20 sec334 33 6

SEC338 340 I342 3 44 34 6 3 48 3 50 35 2 3 54 35 6 35 8

Time from L i f t o f f (sec)360

Pig. H-22. GE Mod III-G Flight Data from SECO to SECO + 20 Seconds

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, ^B iCONFIDENTIAL

-400-500-600- 7 0 0-800-900

-1000-1100-1200-1300-1400

-1500-1600-1700-1800-1900-2000

Predicted nominal wind run 8 Q - G T - 6 (final)*GE Mod III- G final flight data*

*IncludesR a w i n s o n d e balloon dataCape K e n n e d y0739 EST, 15 December 1965

\ \\ \\ \

-SECO+ 20 (358.737 sec)

'

100

Predicted SECO + 20 (356. 894 sec)-120 140 160 180 200 220 240 260 280 300 320

Time from L i f t o f f (sec)

340 360

Fig. 11-23- Cross-Bange Velocity (Yp) Versus 1

' -

NFIDENTIER 13227-6


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200

100

- 1 0 0

.-

-200

-300

-400

-500

-600

-700

- 34 L

SECO + 20 (358.737 sec)

Predicted nominal wind run 80-GT- 6 (final)*GE Mod III-G final flight data*

100

IncludesRawinsonde balloon dataCape Kennedy0739 EST, 15 December 1965

Predicted SECO + 20 (356. 894 sec)

;

120 140 160 180 200 220Time from Liftoff (sec)

240 260 280 300 320 340 360

Fig. II-2k. YawSteering Velocity (VY) Versus Time

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- 35

110

100 -


~ 60 -< L >

40

20 40W i n d Speed (kn)

- 300 280 260 240W i n d A z i m u t h (deg f r o m n o r t h )

Fig. 11-25. Wind Sp eed a n d A z i mut h V e r s u s Altitude

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I I - 3 6

110

100


I Temperature

Pressure (psi)-80 -60 -40 -20

Temperature ( C )Fig. 11-26. A m b i e n t Temperature and Pressure Versus Altitude

+20

ER 1 3 2 2 7 - 6


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^

FDNA


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COE1376


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7 . l l a s i m u m D yn am ic P r e s s u r e- - - -T h e m a x i m u m d y na m ic p r e s s u r e f o r t he G T - 6 t r a j e c t o r y \vas l e s st h a n d e s i g n l i m i t s . T a b l e 11-9 c o m p a r e s t he p r e d i c te d a n d o b s e r v e d

c o nd i ti o ns a s s o c i a t e d w i th t h e m a x i m u m d y n a m i c p r e s s u r e . The p r e -d o m i n a n t l y ta i l -m i n d e n v i r o n m e n t f o r t h i s f l ig h t i n i t s e l f r e d u c e s t h em a x i m u m d y na mi c p r e s s u r e . A p r e d i c t e d t r a j e c t o r y c o m p u t a ti o n f o ra no wind condi t ion showed tha t the max im um dynam ic p r es su re lvouldbe 737 . 5 p s i , a n d t h e p r e d i c t e d t r a j e c t o r y w it h T - 1 w i nd s , f r o m T a b le11-9, sho ws a va lue of 729. 7 ps i , ver i fying the effec t of a ta i l wind.H o w e v e r , t h e o b s e r v e d m a x i m u m d y n am i c p r e s s u r e w a s t h e s a m e a st h e p r e d i c t e d n o win d v a l u e . T h e r e f o r e , o t h e r f a c t o r s su c h as enginep e r f o r m a n c e a n d t he TARS p i tc h p r o g r a m m i n g c o m bi n e d to i n c r e a s et h e m a x i m u m d y n am i c p r e s s u r e s l i g ht l y c o m p a r e d t o t h e p re d i ct i on .

T A B L E 11-9T r a j e c t o r y P a r a m e t e r s a t M a xi m um Dy n am i c P r e s s u r e

I D y n am i c p r e s s u r e ( p sf )T i m e f r o m l if to ff ( s e c )Predicted:k(nomina l ) Obse rved**

Rela t ive wind veloci ty ( fps)Wind veloci ty ( fps)

I

Ma c h n u m b e rAlti tude (ft)Rela t ive f l ig ht pa th angl e (deg)

I W in d a z i m u t h ( d e g f r o m n o r t h ) 2 81Angle of a t t ack (deg) 0 .611.8947, 65046.28

Angle of s id es l ip (deg)*Ref. 10, u pdated ( see footnote to Tab le 11-7)**Mod 111-G 10 pps r ad a r da ta

0 . 8 6 -0 .24


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I I - 40

8. Angles of Attack and SideslipP red icted and observed histories of angles of attack and sideslip dur-

ing the ascent are shown in Figs. 11-12, 11-13, 11-20 and 11- 21. Thepredicted values were obtained from a digital run utilizing wind and atmo-sph eric infor m at ion obtained fro m the 0739 EST Rawinsonde sounding.Observed angles of attack and sideslip were derived using the Mod III-Gposition and velocity information, IG S attitude data and the aforemen-tioned weather data.

. PAYLOAD CAPABILITYP rop ellants rem aining onboa rd afte r Stage II low level sensor un-cover indicat ed tha t a bu rning tim e m argin (BTM) of 2. 348 seconds

existed to a com ma nd shut do wn. The tot al pro pellant weight ma rginwas 767 pounds, and the corresponding G LV payload capability was8655 pounds. These values and the predicted nominal and minimumvalues appear in Fig. 11- 29. The predicted capability curves weretaken from the G LV- 6 preflight report (Ref. 13), updated to incorporatethe 81.4- de gree launch azim ut h, yaw steering to correct for the 0. 2002-degree wedge angle, revised guidance consta nts, and the - 1.07% pitchand - 1 .4% roll programmer biases. The predicted propellant weightand burning time margins are based on the differe nce between thesecurves and the 7821- pou nd spacecraft weight.Real- time payload predictions differed from the predictions shown

in Fig. 11- 29 because extr apo late d actual propellant tem pe rat ur es wereused instead of preflight p redicted propellant t em pera tu res. The lastpayload prediction indicated that th e minimu m payload c ap ability was275 pounds more than the spacecraft weight, and the nominal payloadca pa bility was 878 pounds greater than the spac ecr aft weight at the pre-dicted launch tim e. The act ua l (postflight re co nstr u ct ed ) G LV capa bilitywas 834 po und s greater than the spac ecr aft weight.

C. STAGINGThe staging sequence was normal and physical stage separation oc-

curred as planned. The time interval from staging signal (87FS 9/ 91FS1)to start of Stage I I engine chamber pressure (P ) rise was 0. 667 sec-C3ond. This co m pa res'favora bly with the nominal expected time of 0. 70*0.08 second. Stage separation occurred 0.015 second following startof P rise.

C3

ER 13227- 6


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>ENTIAL - 41

D. WEIGHT STATEMENTTable 11-10 shows the G T- 6 weight history from launch to orbitalinsertion.The postflight weight report (Ref. 11) provides the background datafor this summary. The report includes a list of dry weight emptych anges a t ETR and shows a derivation of weight em pt y fro m th e a ctu alvehicle weighing. Other i tems covered include the derivation of burn-out, BECO, SECO and shutd own weight s; weight co m par isons with theBLH data; and the center of gravity travel envelope as a function ofburn time for the horizontal, vertical and lateral planes.

TABLE 11-10G T - 6 Weight Summary

Loaded weightStart and grain lossesTrajectory LO weightPropellant consumedto BECOCoolant waterFuel bleedWeight at BECOShutdown propellantStage I burnoutStage II engine startGrain lossStage II LOPropellant consumedto SECOAblative, covers andcoolant waterStage II at SECO

Weight (Ib)Step I

272,9253,617(3)

269,308257, 533

0

11, 775164

11,611(2)11,611

Step II65,343

65,343

1165,332

65, 332188

365,14158,754

206,367

Step III7, 821

7 , 821

7,821

7,821

7,821

7,821

Stage Total346,089

342,472

84,928

84,764

72, 962

14,188(4)

CONFIDENTIALER 13227- 6


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II-42

TABLE II-10 (continued)

Shutdown propellantWeight at SECO + 20seconds

Weight (Ib)Step I Step II

13 66 , 2 3 1

Step I I I

7,821

S ta g e Total

14,052 ( 4 )

( 1) Information f r o m N A S A - H o u s t o n(2) Includes outage: 834-lb Stage I; 255-lb Stage II(3 ) Event: launch bolts blown(4) Includes 76 7 I b o f usable propellant

ER 13227-6


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II-43

G T- 6A Flight Test Values

L

Spacecraft weight = 7821 I b

Y, Phase pane / ' / 20 40 80 100

Time i n L a u n c h W i n d o w ( n i i n )Fig. 11-29. Payload Capability

Targe t ing ;change120 1 4 0

E R 1 3 2 2 7 - 6


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- 1

III. PROPULSION SYSTEMA. LAUNCH ATTEMPT ( 12 DECEMBER 1965)

Th e launch of GLV-6 on 12 December 1965 was automatically ter-minated at 87FS + 1. 158 seconds due to premature separation of theelectrical tail plug disconnect 3D1M.

GLV-6 Stage I engine experienced a normal start transient throughthru st cham ber ignition and into secondary rise until start cartridgeburnout and gas generator ignition. S/ A 1 s ta r t transient appearednominal and nearly reached equilibrium conditions at the time the shut-down comm and was initiated. S/A 2 thru st chamber pressure rise wasnormal th rough the completion of start cartridge burning, but the sub-assembly failed to achieve satisfacto ry bo otstr app ing and all parameterss ta r ted to decay abnormally at 87FS. + 1. 05 seconds. S/A 2 datashowed that sufficient energy was generated by the start cartridge toprovide proper bootstr apping operat ion. Gas generator c ombust ion inS/ A 2 apparently was not sustained, and a restric tion in propellantflow to the gas generator was the suspect ed cause.

To determine th e exact cause of engine perfo rma nce decay, t he S/A2 gas generator, fuel and oxidizer bootstrap lines, fuel and oxidizercheck valves and s t ra in ers were removed for inspection. A smal lplastic dust cap was found lodged in the gas generator oxidizer injectorinlet, which prevented oxidizer flow to the gas generat or.

As sho wn in Figs. I I I - l and III -2, both S/A 1 MDTCPS and S/A 2MDTCPS actuated and de-actu ated during the flight attempt. S/A 1chamber p r e s s u r e exceeded the TCPS actuation tolerance band (600 to640 psia); there fore, it is conc luded that S/A 1 TCPS had actu ated .As shown in Fig. III- l, S/A 2 chamber pressure reached th e TCPSactuation limits only mom entarily. Both subassem bly switch es mustactua te before TCPS make signal is given, and because this signal wasnot given for GLV-6 a t t e m pt , it was assumed that S/A 2 TCPS did notmake.It was concluded that, had the disconnect malfunction not initiatedshutdown, the engine would have received a shutdown c o m m a nd a t +2. 2 seconds due to TCPS not being actuated.The fuel autogenous system was functioning normally, which is in-dicated by the fuel pressuran t differential pressure switch (FP DPS)actuation at 87FS + 0. 98 second as shown in Fig. I I I -3. The oxidizer

pressuran t orifice inlet pressure (POPOJ) also shown in Fig. III- 3 didnot reach the OPPS actu ation l imits of 300 to 445 psia; hence, OPPS didnot actuate. POPOI reached a maximum of only 212 psia.

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' . t - -1000

900

--.':

.

-

800

700

600

500

400

300

200

100

P (Meas 0003)

M D T C P S (Meas 0356)

: *

0. 5Time from 87FS1 (sec)

Fig. III-l. Launch Attempt S/A 1 Start Transient 1

^ BW1*J l~ *


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1000

900

800

700

600

500

400

300

100

M D T C P S ( M e a s 0 3 5 7)

-I J L_

III-3

200

0. 5 1.0 1.5Time f r o m 8 7 F S . (sec)

2. 0 2. 5 3.0Fig. III-2. Launch Attempt S/A2 Start Transient

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1000

800

--

60 0

-I 400

2 0 0 ;

1 1 1

87FS

j^

V S .. MDTCPS (Meas 0356)

_

1I + 0 . 5 + 1.0 + 1. 5

Time from 87FS. (sec)+2.0 +2.5 + 3 . 0

>NFIDENTIAL Fig. IH-k. S/ A1 Start TransientER 13227-6


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I l l -7

P transducer was wrapped with additional thermal insulation toC levaluate the effects of thermal environment on transducer dri ft.

Following the 25 October 1965 countdown and during the time ofGLV-6 bonded s torage , al l prevalves were removed. New prevalveswere installed prior to the 12 December 1965 launch a t t e m pt .b. Start transientTh e S/A 1 and S/A 2 thrust chambe r s ta r t transients were normalas shown in Figs. III- 4 and I I I -5. The ignition spikes indicated 89%of ra ted thru st for both S/ A 1 and S/A 2, which is above the enginemodel specification allowable (75%). However, the Gemini P instru-

mentation has charact eristically shown undamped oscillations whichobscure the true transient performance and prevent accurate deter-mination of the ignition spikes. Significant start events are presentedin Table III- l.

TABLE III- lStage I Engine Star t Parameters

ParameterFS , to initial P rise (sec)1 P ignition spike (psia)P step (psia)P overshoot (psia)

S/ A 10.736692465None

S / A 20. 751696445None

c. Steady- state perform anceStage I engine flight performa nce agreed c losely with the preflightpredict ion. Flight integrated average performance parameters werewithin 1. 0% of t he preflight predicted.Engine performance was calculated from measured flight data withth e Martin-Baltimore P RESTO program and used th e Stage I th rus tcoefficient relationship as modified by Mart in. Th e modification in-creased thru st and specific impu lse app roximat ely 3400 pounds and2. 0 seconds, respectively, above th e values calculated with th e Aerojetth rus t coefficient re lationship. The Martin-modified thrust coefficientalso was used in the preflight pr edictions.

)NFIDE J T I A j f cER 13227-6


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CONFIDENTIAL 1 1

I_--

-- -

L U O U

800

60 0

400

200

0,

NF

'

,

''

"...-..~.

t

f %i !

IDENTIAL 13227 -6

"..~~

* 11V+0.5

I

:~f.Jsi

i

/p*

. J/ *r

/1i f wif;.

.

Aif

.,;

+1.0

.:

tflW*fy*\*sr

i

.

J . . '

^nA*

.....

W4f** Y*V

.

.

P (Meas 0004)C2

lyxV*"4*

-

+ 1.5Tme from 87FS (sec)

7*~MDTC..........

.

nA*%JWf*-

:PS(Mea-

\

1 '

s 0357)

/w

+ 2.0

-WfDENTIAL

vvwvWW

i .

+ 2.5

Fg. III-5

1

WuVrfV... . :

;

- .

.- i

.

w _

'

+ 3.0

S A 2 Start Transent


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Ill-9

The Stage I engine average flight performance, integrated froml i f to f f to 87F S, is compared with the preflight prediction in Table III-2.

TABLE III-2Predicted and Flight Performanc e Compari son--St age I Engine

ParameterThrust, engine (Ib)Specific impulse, engine (sec)Mixture ratio, engineOxidizer f l o w rate,overboard (Ib/sec)Fuel flow rate, overboard(Ib/sec)

Preflight*Average455,7422 7 6 . 851.9248

1 0 8 2 . 97

5 6 3 . 17

Flight*Average453,7932 7 7 . 271 93811079 295 5 7 3 8

D i f f e r e n c e(% )- 0 . 4 3+0. 15+0. 69-0. 34

-1 03*Martm-Baltimore modified thrust coefficient relationship used

Engine perform ance calculated throughout the Stage I flight is pre-sented in Fig. Ill-6. The preflight prediction is also shown forcomparison.The S/A 1 thrust chamber pressure transducer was wrapped withfour times the normal amount of thermal insulatio n. The extra i nsula-tion was used on the transducer to verify that the P transd ucer drift

was due to thermal effects and to confine the drift to acceptable limitsRecons tructed data showed that the P tran sduce r, which had normalC2insulation, began to drift at approximately 87FS + 80 seconds anddrifted approxim ately -2% (normal for S/A 2 as established from pre-vious Gemini flights) P data showed no negative drift, verifying thecltheory that P drift was due to the excessive thermal environment.

Stage I engine flight performance calculated at the 87FS.. + 55 secondtime slice and corrected to standard inlet conditions is shown in TableIII-3. This is compared to the acceptance test and the predicted flightperformance at standard inlet conditions and the nominal time as used

ER 1 3 2 2 7 - 6


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CONFIDENTIAL

|-600 580|2560 >>

aj 540 n)

U(p f.-DO

U1DO

:-aa

1-0)1120

1100108010601040

- ^

- g0)

- 'SoW-^

2.001.951.901.851.801.75

. >U

28528027527 0265260255

i .h

-

490480470460450440430420410

. _ > _ &_0

Q Q Q G

20 10 6 0 I 100 120 140 160Lift-|off8 7 F S ,

T i m e f r o m 87FS. (sec)

A v e r a g e E n g i n e P e r f o r m a n c e I n t e g r a t e df r o m Liftoff to 87FS.

S y m b o lF t(lb) '8"MR eW ( I b / s e c )W fo ( l b / s e c )

Pre f l i g h tP r e d i c t i o n4 5 5 . 7 4 22 7 6 . 8 51 . 9 2 4 81 0 8 2 . 9 7563. 17

F l i g h tA v e r a g e4 5 3 . 7 9 32 7 7 . 2 71.93811 0 7 9 . 2 95 5 7 . 3 8

P r e f l i g h t p r e d i c t i o n F l i g h t p e r f o r m a n c e

Fig. III-. Stage I Engine Flight Performance

CONFIDENTIALER 13227-6


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in the r e f O . he ~ ~ e d i c t elight perJmanCe a t standard'OnditiOns LVaS ob ta ine d by modif yin g the nomi na l acc ep tan ce test dataa 4 8 5 0 - ~ 0 u n d c c ep t an c e - to -fligh t th ru st growth obtained from analv-of PI.evious Titan I1 and G Lv flights.


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III-12 ITIAL

TABLE III-4Stage I Engine Shutdown Parameters

ParameterTime from P decay to 87FS (sec)P at 87FS (psia)Time from FS9 to data dropout (sec)P at data dropout (psia)

S / A 10 . 72000.7150

S / A 20 . 72500.7148

e. Engine malfunction detection system (MDS)The Stage I engine MDSoperated satisfactorily and within specifiedlimits throughout the flight. Figures III-4 and III-5 illustrate responsetimes and actuation levels of the malfunction detection thrust chamberpressure switches (MDTCPS) during engine start for S/A 1 and S/A 2,respectively. Figures III-7 and III-8 show deactuation times and levelsduring shutdown for S/A 1 and S/A 2, respectively.A summary of the operating characteristics of the switches is tabu-lated in Table III-5.

TABLE III-5Stage I MDTCPS Operation

SwitchS/A 1S / A 2

ActuationTime (sec)FS1 + 0. 900FS- + 0. 920

Pressure (psia)585575

DeactuationTime (sec)FS - 0. 047FS - 0. 045

Specification RequirementsActuation 540 to 600 psiaDeactuation 585 to 515 psia

Pressure (psia)550530

ON FIDENER 13227-6


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1000

800

IH-13/5V

i-4001200

P (Meas 0003)t spPgflSM

-2


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m - 1 4l O O O ONFIDENTIAL

(Meas 0004)1

w3 600'Staging "j.blackout .

M DT CP S (M e a ' s 0357)f j ;_ ; *^ * \ivieas UO O""'-' - ..j..."- ' 1 ' ' J :till-1.0

I + 1.0' \Time from 87FS2 (sec) 87FS +

Fig. III-8. S/A 2 Shutdown Transient



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ONFIDENTIAL I I I - 1 5

f. Engine prelaunch malfunction detection system ( P M D S )All PMDS switches actuated within specified actuation times andpressures as shown in Table III-6. As a result of the later than ex-pected OPPS actuati on on GLV-7 launch, the oxidizer pressurant backpressure orifice diameter was c hanged from 0. 50 inch to 0. 46 inch.The smaller back pressure orifice increased the oxidizer pressurantorifice inlet pressure (POPOI); consequently, the OPPS actuated earlierthan on previous Gemini flights and substantially earlier than the inter-rogation time (T-0 + 2 . 2 seconds), as shown in Fig. III-9.

TABLE III-6Stage I PMDS Operation

Actuation timeMeasured time from87FS (sec)Measured time fromTO (sec)Required time (sec)*

Actuation pressureMeasured (psia)Required (psia)

TCPS

0. 981

1. 043T+2. 2#*60 0 to 640

OPPS

1. 5781 . 640T + 2 . 242 436 0 to 445

FPDPS

0 . 9 1 20 . 9 7 4T + 2 . 2**46 to 79 (psid)

*The shutdown'timers start from Tn.**Not instrumented.2. Stage II Engine (YLR91-AJ-7 S /N 2 0 0 7 )

a. Configuration and special proceduresThe GLV-6 Stage II engine configuration was identical to that of GLV-7.During the time of GLV-6 storage, afte r the 25 October 1965 launchattempt, the Stage II gas generator was removed and returned to Aerojet,Sacramento, for cleaning.

E R 1 3 2 2 7 - 6


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- /6 - 1-M l'

a :_1- .iL-1

"1J

I-3.-

i---_.i,

600

1 0

300

150

0

(

-

)x i d i z

.-

I

sr pre SSIr_ irant pre'

(

3SU releas 2105

+ 0.5

.)sw i t c h i ( O P P S )

'

P O P O l 2 ( M e a s

*

I3PPSsresst~a n g e

'

;

"m ak e

0026)

.

+ 1.0

..

II

'

. **

_

.

Ii1111L

'

*

'

+ 1. 5

. ,

1

:

,.

"~~^ OPPS' i n t e ri.

:

rogati:

on

'

.

..

+ 2.0 +2. 5 + 3.8 7 F S , T i m e , f r o m T - 0 (sec)

F i g . III-9. S/A2 Start T ran si e n t



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After the 12 December 1965 launch a ttempt, Stage I1 r e c yc le r e qui r e -ments were based on Aeroje t Engine Test Directive 2, 1-3. 4E. The TPAoil rvas changed as a precautionary measu re . The prevalves remainedinsta lled, and re loading the propellants direc tly on the thrust chambervalves had no adverse effec t on the engine system.. b. Start transient

Stage I1 engine s tart transient was normal, as i l lustra ted by thethrust chamber pressure in Fig. 111-10. Significant engine s tar t eventsa r e pr e s e nte d in Ta ble 111-7.

TABLE 111-7Stage I1 Engine Start Parameters

PC ignition spike (psia)3

Pa ra m e te rFS1 to initia l PC r i s e ( s e c )

3

Flight Performance0. 651

1 I 1:::Staging blackout period.

/ pc 3 overshoot (psia)

c . Steady-sta te performance

Not available::< I

Stage I1 engine s teady-sta te f l ight performance was sa tisfac torythroughout f l ight and agreed c losely with preflight predic tions. The


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CONFIDENTIAL J J T O F f f l E N T I A L1 0 0 0

800

00

j 600 -gor.i-- 400

200

Time from 91FS1 (sec)+ 3 . 0

Fig . 111-10. S/A 3 Start Tr s ms ien tCONFIDENTIALE R 1 3 2 2 7 - 6 C O N F I D E N T I A L


75/253


76/253

0 CONFIDEN1- m i B E N T I A L

" 104

--.

1.85rt(

-.I01.s

01-

-_< :

"F

V "-- +1 _-0r_ -

hr-i

,

xt

1

1

.

:

+ 291FSr Time fro m 91FS? (sec)

Fig. 111-12. S/A 3 Shutdown Transient

CONFIDENTIALER 1 3 227-6


79/253

C O N F I D E N T I A L C fW - 23 ^5 0 0 0

40 0 0

30 0 0

2000 h

1000

1 Fig. 111-13. Stage Engine Thrust Tail-OffCONFIDFMTi A i


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III-2 4 CONFIDENTIAL

switches (MDFJPS) during the start and s h u t d o w n transients, respec-tively. The f u e l injector pressure is not instrume nted and, ther efore,is not available. A summary of the significant switch parameters ispresented in Table III- 10.TABLE III- 10

Stage II MDS OperationParameter

A c t u a t i o n time (sec)P at actuat ion (psia)D e a c t u a t i o n time (sec)P at deactua tion (psia)

91FS. + 0 . 7 3 752 091FS + 0 . 14 045 0

. PR OPELLANT SUBSYSTEM1. Propellant Loading

a. Loading procedureFive propellant loadings were performed on G L V - 6 , consisting ofthe RTP and W M S L exercises, two launch attempts and the actuallaunch. (See Table III -11.)

TABLE I I I - 1 1G L V - 6 Loadings

OperationRTPW M S L1st l a u n c h attempt2n d launch attemptL a u n c h

DescriptionD u a l loadingD u a l loadingD u a l loadingD u a l loadingD u a l loading

Date28 September 19657 October 196524 October 196512 December 196515 December 1965

All loadings were made using the tandem flowmeter system installedafter the l a u n c h of GT- 5. No serious ground or airborne hardware



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prob l em s o r c u r r e d d u r i n g the f ive propellan!. Inadings; h o \ \ c v ~ : - ~ ,! i~nl!3~:rf fluv:rneters \rVere enloved folio\ring the 15.21SL a n d ai;a.iilfollc\~:inethe 1 2 Octobe r 1965 iaunch a t tempt .

Tile check ca l ibra t ion results of the flow meter^ removecl fro111 t h es ~ ~ s t e n :r e pr e s e n t e d i n T a l ~ l e 11- 13. These data h a ~ ~ e1een upplietlto t he i i f f c r e nc e s obse rl -ec l b ~ t~ x . e e nloxvmeter and t a b r .un ~ v h e r c \ . ~ rappl icable .

F low m e t e r V e r i f i ca t i on R e s u l t s

The tab r u n s u s e d fo r the f ive loading opera t ion s lx!ere d er i ve d a sfol lows:

-1'est i \ f ter[Vhich \letel*[['as Checked

IVlISL

WXISL

LVMSL

\YITS -

(1) R T P Obta ined f r om Denver t an k cal ibra t ion da ta .2 W?;ISL Obta ined f ro m D e n v e r t ank ca l ibra t ion da ta .

AIeterPos i t ion

Stzge I fue l

S tage I fue l

S tage I1 fuel

S tage I1 fue l

( 3 ) First Launch Or ig ina l t ab run c o r r e c t e d f o r R T P an d:Ittempt WAISL r e s u l t s i ncluding f low m e te r ve r i f i -c a t i ons a f t e r t he W!lISJ,.

S t a g e I I f u e l

l l e t e rN o .

2 0 2 146 3Iarti11-Denver D e n v e r109 1 7 2 Mart in- XIar t in -

\ l a r in -

206361

20636 1

Denver 1 D e n v e rDenver

Denverh l a r t i n -DenverM a r t i n --

S t a g e I f u e l1st !at:nch Denvera t t e m pt i Stag e I1 fuel 199170 / Ma r t in -II j Denver

Denx~erf i la r t in-DenverLf'yle- -Lla r t in- 1 +0. 1

+O . 3 9

+O. 14

DenverMa r t i n -Denver [email protected]


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Ill-26

Stage IFuelN o change

Stage IIFuel-0. 31%

Stage IOxidizer+ 0 . 14%

Stage IIOxidizerN o change

( 4 ) Second Launchattempt Tab runs revised to account for results offirst three tests, flowmeter verificationsafter the first launch attempt and the d i f f e r -ence between Denver and Wyle calibrationfacilities. Changes below are from originaltab run.

Stage IFuel+0 . 0 4%

Stage IIFuel-0. 30%

Stage IOxidizer-0. 11%

Stage IIOxidizer- 0 . 2 3 %

( 5 ) Launch Tab run pounds changed to account for openprevalves; otherwise same as that for sec-on d launch attempt.Tests at the Denver and Wyle calibration facilities have establishedthat, if a f u e l or oxidizer flowmeter calibrated at Martin-Denver isassumed to be correct, a corresponding Wyle meter will read about0. 3% higher. It is not k n o w n which facility is more nearly correct;however, the launch loading was based on the Martin-Denver calibratedflowmeters being correct. This, in e f f e c t , decreased the Wyle cali-brated flowmeter/tab run errors by 0. 3% and established the leastprobability of payload loss.A detailed summary of results of the f ive propellant loadings madefor GLV-6 is shown in Table III 13.

E R 1 3 2 2 7 - 6


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NFIDENTIAL - 27

TABLE III-13Summary of Propel lant Load Verification

L o a d i n gEventRTF

WMSL

FirstL a u n c hattempt

SecondL a u n c hattempt

Launch

TankStage I f u e lStage II fuelStage I oxidizerStage II oxidizer

Stage I fuelStage II fuelStage 1 oxidizerStage II oxidizer

Stage I fuelStage II fuelStage I oxidizerStage II oxidizer

Stage I f u e lStage II fuelStage I oxidizerStage II oxidizer

Stage I fuelStage II fuelStage I oxidizerStage II oxidizer

Flowmeter

SerialNo.2021941991722063611991702042772063592061681991732021461991722063611991702 0 4 2 7 7206359199168199173199169202146199171199170SPOOL202164199168199173206362202146206361204278206360202164199168199167206362202146206361204278206360202164199168199167

FTPSNo.441

441144544194424412446

44204414411445

4419442

4412446442044144114454419442

44124464420441441144544194424412446442044144114454 4 1 944244124464420

ConnectedtoCP 2500Counters

XXX

XX

XX

X

XX

XX

XXXX

XXXX

Calibra-tion Facil -ity (1)DDDDWWWWDDDDWWWWDDDD-WWWWDWDWWWWWDWDWWWW

H i - LiteTemper -a tu r e (*F)44.044.046.946.948.648.649.849.829.929.930.830.826.926.929.429.434.534.533.833.836.236.237. 537.528.228.229.029.026.626.627.627.629.229.230.030.029.429.429.829.8

DifferenceBetweenFlowmeteran d ActualTab Run Nom(% ) (2 )+ 0 . 141+0.398+ 0.086- 0.070+ 0.168+ 0.318+ 0. I ll+ 0.152+ 0.358+ 0.645+ 0.174+ 0.008+ 0.082+ 0.084+ 0.032- 0.109+0. 584+ 0.006+ 0.065+ 0.170- -+0. 153+ 0.147+ 0.081+ 0.469- 0.022+ 0.477+ 0.243+ 0.274+ 0.378+ 0.355+ 0. 113+ 0.497- 0.001+ 0. 516+ 0.268+ 0.288+ 0.420+0.364+ 0.183

DifferneceBetweenFlowmelerand FirstTab Run(% ) (3 )- 0.025- 0. 192- 0.304- 0.340+ 0.168+ 0.318+ 0. I ll+ 0.152+ 0. 192+ 0.065- 0.216- 0.262+ 0.082+ 0.084+ 0.032- 0.109+ 0.174+ 0.006- 0.245- 0.410- -+ 0.293+ 0. 147+ 0.081+ 0.509+ 0.018+ 0 .1 7 7- 0.057+ 0.164+ 0.268+ 0. 125- 0. 117+ 0.537+ 0.039+ 0.216- 0.032+ 0.178+ 0.310+ 0. 134- 0.047

Allow-ableToler-ance(%)

t o .

tO. 110.31*0. 1

AverageLoad- inFlow Rate(gpm)2 4 312 1205

97

23010917 585

(1 ) W Wyle laboratories D = Martin- Denver(2) Actual difference observed during loading, no t corrected for meter verification where applicable.(3) Difference from RTF or WMSL tab run corrected for f low ra te and meter verification resul ts whereapplicable. (Does no t include any meter bias.)



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Ill- 2 8

The sequence of pro pellant loading events is given in Table III - 14.

TABLE I I I -14G T ~ 6 Launch Propellant Loading Schedule

EventStart pre chillStart loadHi- liteLoad complete

Time (EST) 14 Dec, 1965Stage IOxidizer

19352012

21422155

Stage IIOxidizer19352013

20502108

Stage IFuel2205223023252333

Stage IIFuel2205223023022307

Mission loads for the oxidizer tanks were obtained by using theK- fac tor ratio technique. This was in accord with a Martin Company/SSD agreement that an oxidizer flo wm e t e r / t a b run e r r o r of more than+ 0. 1% at hi- lite would constitute an out ~of- tolerance condition.A flowmete r - t o - t ab run comparison is shown in Figs. I l l- 14 and1 - 15. In each figure, the data are referenced to the tank calibration

made at Denver (which is synonymous to the special loading tab run).The data fo r Wyle calibrated m e t e r s are not corrected for the differ-ence between Denver and Wyle facilities. In Fig. Ill- 14 the applicationo f the - 0. 3% correction to all Wyle resu l t s will account for the place-ment of the launch tab shift.b. Total propellant loadsTotal mission loads for the launch, as det erm ined from flowm ete rs,are shown in Table III - 15. The flowm et er tot alizer r ea dings wer e c or-rected by subtr ac ting prop ellant vapo rized and propellant remaining inthe fill lines. Oxidizer flowmeter loads reflect the use of the K- fac to rratio method to obtain m ission load s. Total pr op ellant loads as deter -mined by flight verification are also shown in Table III- 15. The flightverification loads were calculated from a pr op ellant invento ry, u singactual level sensor uncover t imes and tank calibration data to determineflow ra tes . Tot al, integra ted , in- flight overboard prop ellant con-sumption was f o u n d using the engine analytical model. Engine starttra nsient co nsum pt ions were der ived from Aero jet sum m ary repor t s .Other t ransient propellant consumptions and pressurization ga s weightswere calcu lated from flight d at a (T ables - 36 and - 37).



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CONFIDENTIAL 111- 29

Data are correc ted fo r flow rate and me ter verification resul ts , whereavailable. All data are referenced to original tank calibration and representthe percent error of the flowmeter resul t at hi- lite from the original calibration.

Stage I- r 0.5

-- 0.4

202164 L - r 0.3

- - 0 .2206360 L

Launch ,t ab shif t '

- - 0 . 1

206359 RTF202164 LA 1202164 LA 2204277 RT F

2 06 36 0 L A 2- 206359 WMSL 2 04 27 7 WMSL

- - 0

r v- - - o . 1

- - - 0 . 2

- "- - . 3Note: All meters are Wyle calibrated

LEGENDR T F = first loadingWMSL = w e t mockLA 1 =25 October 1965 launch attemptLA 2 = 12 December 1965 launch attemptL = launch

Stage II- 0.5

-- 0.4

-- 0.3

0. 2 199168 L-

199167 L

Launch , r ^t ab shift 1- u '

- - 0. 1

199173 RTF199168 LA 1199168 LA 2206168 RT F199173 LA 1

199168 WMSL- - 0

0. 1 199173 WMSL199167 LA 2

- 0.2

-"--0.3

Fig. III-lA. GLV- Loading SummaryOxidizer

CONFIDENTIALER 13227 - 6


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111-30 CONFIDENTIALData are correc ted fo r flow r a te and met e r verification results, whereavailable. All data are referenced to original tank calibration and representthe percent error of the flowmeter result at hi- lite from the original calibration.

W 202362 L

Stage I- i- O. 6

05*- W 202362 LA 2

-0. 4

- -0 .3

0. 2- = ,

D 202146 LLaunch ,- -tab shift ' '

- - 0 . 1D 202146 WMSLD 199169 L A 1D 199172 WMSLD 202146 LA 2D 202146 LA 1D 202146 RTF0. 1

_ 0 2_ r D 199172 RT F

0. 3

0.4

-"-- . 5

W 206361 L -

Stage II- T - 0 . 6

--0. 5

- -0. 4

--0. 3

2

- - 0 . 1

D 204278 L

- 0. 2-jLaunch , Jt ab shift L i

-0.4--

W 206361 L A 2

D 204278 L A 20.1

-0. 3

D 206361 WMSLD 199171 L A 1D 199170 WMSLD 206361 RTFD 199170 RTFD 199170 LA 1

- L-0. 5

LEGENDD = DenverW = WyleRTF = first loadingWMSL = w e t mockLA 1 = 25 October 1965 launch attemptLA 2 = 12 December 1965 launch at temptL = launch

Fig. 111-15. GLV-6 Loading Summary--Fuel

CONFIDENTIALE R 1 3 2 2 7 - 6


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p

3r50|>Q08005>4~)CU01_Q3)~

,

->*4

8~&

rf(0)_g-:-(1.id-CODNA

E16


88/253

I l l- 3 2

. P r o p e l l a n t assayP r e l a u n c h d a t a f r o m t h e p r o p e l l a n t assay repor t ( s a m p l e d o n

6 D e c e m b e r 1 9 6 5 ) fo r o x i d i z e r a n d f u e l a r e presented i n Table I I I - 1 6 .S p e c i f i c a t i o n values are also listed. G o o d agreement was o b t a i n e db e t w e e n t h e analysis a n d s p e c i f i c a t i o n requirements. D a t a a r e f r o mt h e pr imary R S V p r o p e l l a n t s w h i c h were u s e d t o load t h e v e h i c l e .

T A B L E I I I - 1 6P r o p e l l a n t Assay S u m m a r y

F u e l A I I L - P - 2 7 4 0 2 ( U S A F )H y d r a z i n eU D M HH9OT o t a l X0H4 + U D M HS o l i d sPart ic les o n 5 0 mesh screenD e n si t y ( g m / c c ) at 77 FOxidizer M I L - P - 2 6 5 3 9 ( U S A F )N i t ro g e n t e t r o x i d e ( N 9 O . )C h l o r i d e as N O C 1H 9 O e q u i v a l e n tSolidsN o n v o l a t i le ashPart ic les o n 5 0 mesh screen

Test5 1 . 7 %47 . 9%0 . 4 %9 9 . 6 %0 . 2 m g / l i t e r00 . 8 9 9 9Test9 9 . 8 %

*0 . 0 1 %

00

R e q u i r e m e n t51 . 0 . 9%4 6 . 9 % m i n2. 0% max98% min25 m g / l it e r0

- -Requi rement99. 4% min

- -0 . 2 %1 0 m g / l it e r

- -0

; N o t r e p o r t e d .

ER 13227- 6


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Ill-33

2. Propellant Temperaturea. WeatherA comparison of the F-45 day prediction, the F~l day predictionand the actu al weather for the 15 December launch of GT~6 is presentedin Table III- 17. The F-45 day prediction was based on weather for ahot December through M arch day. There was, in general, better agree-ment between the F-45 day prediction and actual weather than betweenbetween the F- l day prediction and actual.Predicted wind speed average was approximately 52%higher thanactual.

TABLE III- 17Predicted and Actu al Weathe r Conditions for G T - 6 Launch

Time(est)2100 02200 Q2300 200000100020003000400 0500 0600 0700 Q0800 -0900100011001200

Dr y BulbTemperature (F)F-457 1 . 270. 770. 169. 869. 569.369. 068.868. 768.568.470. 972. 875.076.477.2

F- l6 7 . 0

63.0

59.0

59.0

63.0

68. 0

A c t u a l68.068. 268.468.868. 569. 069.370.569,268. 467. 367. 370.1

Dew PointTemperature ( F)F-4565.865.364. 964. 664. 464,264. 063. 963. 763. 664. 966. 167.266.268. 869.2

F- l61. 0

58.0

53.0

55. 0

57. 0

60. 0

Actual66. 066.066. 066. 066.066. 066.066.066. 065.064.064.066. 0

W i n d Speed (kn)F- 45

7777777777779

101212

F- l7

6

6

6

9

13

Actual3144746776567

Cloud CoverF- 450. 50.50.50. 40. 50.50. 50. 50.50.50. 60. 60. 60. 60. 60. 6

F- l0. 2

0.2

0. 2

0. 3

0. 4

0.4

Actual0.80. 61.01.00.80. 80. 60.80. 80.60. 60. 50. 4

b. Propellant loading temperatures

Table 111-18 compares the requested propellant temperatures at theRSV (at start of loading) and the tank bottom probe (at hi- lite) with themeasured propellant temperatures.

ER 13227- 6


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- 34

TABLE - 18Propel lant Tem perat ure Com parison- -RSV and Ta nk Bottom Pro be

SystemStage I fuelStage II fuelStage I oxidizerStage II oxidizer

RSV Tempera tu r e ( F)Requested26. 726. 725.525. 5

Actual26 .526.525.525. 5

0.20.2

00

Tank Bottom ProbeTempera tu r e ( F)Requested28.428.527 . 929 . 7

Actual28.228. 529. 828.8

- 0. 2- i . o+ 1.9- 0. 9

The requested oxidizer RSV tem perat ures were m atched exactly,and the f u e l RSV and fuel and oxidizer tank bottom probe readings werewithin an acceptable range of accuracy.RS V and flowm ete r tem per atu res reco rded du ring loading are shownin Figs. Il l- 16 and III -17.c. L i f t o f f t empe ra tu r e sA comparison of predicted, actual and reconstruc ted propel lantbulk temperatures appears in Table III- 19.

TABLE II I- 19Pro pellant Bulk Tem pera tur e Comparison

SystemStage I fuelStage II fuelStage I oxidizerStage II oxidizer

F-45 DayPrediction(F)38.737.839.842.4

F- l DayPrediction(F)38. 138.239.843.6

Actual(F)41.141.842.044.2

Reconst ruc ted(F)40.839.241.644.0

irrr m inER 13227- 6


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--'i

I

(

aue

CODNA

E1376


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111-36 - " ' U u i i , i l l ' i 1 1 T *

Time of event - Jtage I load complete

Stage I fuel Hi- LiteStage II load comple te

- ! Stage I I fuel Hi- LiteResume loadStart leak checkStart loading

3 5 ;

30

:Meas 4432'(Stage ' f lowmeter)

Meas 4431(Stage I)f lowmeter)2100 2130 2200 2230 2300 2330

Eastern Standard Time (hr)Fig. III-1T. Fuel Temperature During Loading

ER 13227- 6


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- 37

Actual bulk tempera tures a t l i f t o f f were obtained from a computerprogram analysis of flight data. The positions of the recons t ruc ted ,actual and pred ic ted t empera tu re s in the mixture ratio band are shownin Figs. I l l- 18 and I I I -19.

Figures - 20 through 111-23 show a comparison of the F- l day t e m -perature prediction, the recons t ruc ted t empera tu re and actual propel-lant probe temperatures during the countdown fo r each pro pellant tank.Correlation of actual and r e c o n st ru c t e d t e m p e ra t u re s was excellent.The difference between the F- l day prediction was due to the differencesbetween predicted and act ual weat her conditions. Avera ge a ct ual drybulb and wet bulb t em p er a t u res wer e 5. 5 F and 8. 3 F higher , respec -tively, than pr edicte d. P redicted winds averaged 3 knots higher .The m aximu m de viation between actual and recons t ruc ted t empera -t u r e s (10. 3% of total rise) occur red in the Stage fuel tank. All otherswere 3. 8% or less.d. Suction t e m p e ra t u re sThe actual pump inlet tem per atu res were in good agreement withthe pred icted tem per atu re pr ofiles. The se data are shown in Figs.111-24 th ro u gh 111- 27. The t r ends of the actual tem perat ure cur veswere in good agreement with those predicted. Deviations may be as-cribed to differences in predicted and actual weather and the differ-ences between optimum and T~0 t empera tu re s . In Table 111- 20 a com-

pa rison is made be t wee n the suc tion and tank bottom pro bes at variou st imes after FS,T ABLE - 20

Prope l lant Temp era tu re Com par i son- - TankBottom Probe and P u m p Inlet

SystemStage I f u e lStage II fuelStage I oxidizerStage II oxidizer

Time(sec)FS 1 + 5FS. + 25FS 1 + 6FS. + 22

Suction ProbeTempera tu re(F)40. 239. 139. 942. 8

Tank BottomT e m p e ra t u reProbe(F)40. 938.339. 043. 3

DeltaTempera ture(F)- 0. 7+ 1. 4+ 0. 9- 0. 5

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111-38

. - - ;.-.

~

50

45

40

F-45 days predictedlaunch window

Reconstructed

MR (optimum)

2 !

-,

;, ; . : : , ' : . ^5 30 35 40 45 i>0 5Bulk Fuel Temperature ( F)

Fig. III-18. Propellent Bulk Temperatures at Liftoff, Stage I

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_Li H IIII il M l ITT - 3 9

.V- . - ...

i

F- 45 days predicted__launch \ y i n d q w

F- l day pre

: : ;: - :MR (maximum)

M R (optimum)iiM R (minimum) '

35 40 45Bulk Fuel Temperature ( F)

Fig. 111-19. Propellant Bulk Temperatures at Liftoff, Stage II

.CONFIDENTSER 13227- 6


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_const

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i

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i- -

-

- F - l day-45,

Fday

ActualReconstructedF- l day prediction

" SP.

'0 8 0 0

111-41

Eastern S t a n d a r d TimeFig . Ill-21. Stage I F u e l Tank Bottom Probe T em pera tu re (Meas



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111- 42

-i.--.

- F - l dayFday

- . I ' ^15

3 0

ActualReconstructedF- l day prediction

2000 2 4 0 0 0 4 0 0Eastern Standard Time (hr)

Fig. 111-22. Stage II Oxidlzer Tank Bottom Probe Temperature (Meas

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3NFIDENTIAL 111-43

F - l day 1 0

45 ^

10

F day-

a 3 5:

0 0ActualF- l day prediction

' *J< ^ - < - -' ^^^"

00 24 0 0 0 4 0 0 0 8 0 0

- ""

E a s t e r n S t a n d a r d T i m e (hr)F ig . 111-23. Stage II F u e l Tank Bottom P r o b e T e m p e r a t u r e (Meas l|60l)

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111-44

-_

I-45

:

;

' i

Q QActual (Meas 0023) Actual (Meas 0 0 2 4 )

F - 4 5 day predictionA Tank bottom probe

4i< b G G G I D 0

,-,U J

G

Q

?

0

1

087FS, ,:1 100

Time from 87 FS, (sec)120 140 160

Fig. 1-24. Stage I Oxidizer Pump Inlet Temperature (Meas 0023 and 0024)

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1

ActualF-45 day predictionA Tank bottom probe

IF - 50v-

-v 4 5H

0 e

20 40 60 80 100 140 IBO87FS1 Time from 87FS1 (sec)Fig. HI- 25. Stage I Fuel Pump Inlet Temperature (Meas 0013)

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- 48

3. Propellant Feed Systema. Feedline transientsThe maximum transient pre s s u re s record ed at the pum p inlet in-strumentation bosses are listed in Table - 21.

TABLE - 21Maximum Transient Pressures at P u m p Inlet

MeasurementStage I oxidizer

(0017)Stage I fuel(0014)Stage II oxidizer(0510)Stage II fuel

(0507)

AtPrevalveOpeningNo dataNo dataNo dataNo data

AtInitialPressureWaveNegligibleNegligibleNegligibleNegligible

AtIgnition(psia)124

Negligible**

AtTCVClosing(psia)NegligibleNegligible

68Negligible

DesignOperatingPressure(psia)215

55260

80*Not available due to telem et ry staging blacko ut .No data were available on the prevalve opening pressure transients,since these valves were opened during the 12 December 1965 launchat tempt and were not replaced. Ignition transient pr e s su r e s were, ingeneral, similar to those of G LV- 5 and G LV- 7 flights. Telemetryblackout norm ally exper ienced d ur ing Stage ignition eliminates dataon sustainer engine ignition transients.

b. P u m p inlet suction pr e s su r e sStage I and Stag3 II static suction pressures at the suction measure-ment boss locations are shown in Figs. IH- 28 through 111- 31, which

present t he pre flight pr edict ed, po stflight rec onst ru ct ed and best esti-mate of actual flight pressures. The postflight reconstructed curveswere based on flight measured values of ullage ga s pressure , axialload facto rs , prop el lant tem per atu res and pro pellant loa dings.The Stage I oxidizer best estimate curve of the static suction pres -sures at the measu rement boss (Meas 0017) co nsists of an average ofthe measured pressure and the two oxidizer standpipe pressures

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111-49

11 0

100

. 80

70

Preflight prediction Postflight reconstruction

Best estimate o f flight suction pre ssu re

87FS, 100Time from 87FS. (sec )120 140 160

Fig. 111-28. Stage I Oxidizer Suction Pressure (Meas 0017)

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IH-50

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1

- Postf light reconstruction Best estimate of flightsuction pressu re

-

- " \\ i ' ;

i!

1 20 40 60 80 100 120 140 160- 1 Time fro m 87FS. (sec)

Fig. HI-29. Stage I Fuel Suction Pressure (Meas OOlU)

CONFIDENTIALE R 13227- 6


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.CONFIDENTIAL IH- 59Table - 25 lists the level sensor volumes and delta volumes usedin the level sensor f l o w rate analysis. The Stages I and II oxidizer andfuel high- level sensor volumes were reconstructed to reflect the volumeswhich were de ter m ined by calibration at Cape Kennedy using the propel-

lant tr ansfer and pr essu rizat ion system. The Stages I and II outagesand shutdown level sensor volumes were calculated using the actualcounts of flowmeter pulses obtained du ring the special loading and theWMSL exercises.TABLE 111- 25

Averaged Volumes at Level Sensor Locations

TankStage I oxidizer

Stage I fuel

Stage II oxidizer

Stage II fuel

SensorHi- levelOutageHi- levelOutageHi- levelShutdownHi- levelOutage

Averaged Volumes(stretch included )(ft 3)

1708.2037. 85

1402. 5 465. 80

285. 5122.41

350.0819.08

Volume( f t 3 )

1670. 3 5

1336. 74

2 6 3 . 1 0

331.00 . Flow r a t e sTable 111-26 presents the pre dicted and the actual volum etr ic flowr a t e s between level sensors.

TABLE III -26P rop ellant Volu m et ric Flow Rate

TankStage I oxidizerStage I fuelStage II oxidizerStage II fuel

Predicted(ft 3 / s ec )11. 7519.8212.2692.087

Actual(ft 3 / sec)11.7349.7352.2652.071



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I I I - 6 0

d. Mixture ratioTable 111- 27 compares the Stages I and II predicted and actual engine

mixture ratios (between level sensors) for the G LV- 6 flight.TABLE 111-27

Engine Mixture Ratio

SystemStage IStage II

Predicted MixtureRatio1.92801.7470

Mixture RatioActual1.94181. 7578

Sensitivity coefficients applied to the delta between the predictedand actual variations in average suction pressure and t emperature be -tween sensor uncoverings yield the information shown in Table II I- 28.TABLE III-28

Mixture Ratio Pressure and Temperature

SystemStage I oxidizerStage I fuel

Pres -sure(psi)- 0. 5

+ 2. 5Total Stage I

Stage II oxidizerStage II fuel

- i . o-1 .5

Total Stage II

MixtureRatio(Pressure)- 0.000810- 0.006250- 0.007060- 0.004200+ 0.005310+ 0.002760

Tempera-t u re(oF)

+ 0.9+ 1.0

- -+ 1.4+ 3 . 2

- -

MixtureRatio(temp)- 0.002086+ 0.001627- 0.000459- 0.003661+ 0. 005325+ 0.001664

MixtureRatio(total)- 0.002514- 0.003028- 0.007519- 0.006002+ 0.008277+ 0.004424

By applying the delta mixture ratio (to ta l) shown in Table 111-28 tothe pre dicted (F- 45 Day) between- sensor m ixtu re ratios, the run- to -run variation can be calculat ed. The m ixtu re ratio deviation along withthe allowable ru n- to - ru n disper sions are sho wn in Table 111-29.

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I I I - 6 1

T A B L E I I I -29Mixture R a t i o Deviation

SystemStage IStage II

Predicted M i x t u r e R a t i o(corrected for pressurea n

Launch Vehicle No. 6 Flight Evaluationx

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