Launch Vehicle No. 8 Flight Evaluation

8/8/2019 Launch Vehicle No. 8 Flight Evaluation

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( N A S A - C 9) L A U N C H V E H I C L E NO. 8 FLI GHTE V A L U A T I O N (C.artin Co.) 290 p N75-75421'

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P R E P A R E D IY

(ACCESSION NUMBER) (THRU)

CR OR TMX OR AD NUMBER) (CATEGORY)

AUNCHVEHICLE NO. 8

FLIGHTEVALUATION (U)

U . S . GovEngineer!Issued as Supplement^to: Gemini ProgramGemini VIIIMSC-G-R-66-U

y: Gemini VIII Mission Evaluation TeamNational Aeronautics and Space AdministratioManned Spacecraft CenterHouston, Texas A p r i l 1966

Los Angeles, California


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E R 13227-8 A p r i l 1966N A S A - M S C - G - R - 6 6 - USupplemental Report 2

L A U N C HV E H I C L E

LAUNCH VEHICLE NO. 8FLIGHT

EVALUATION (U)

Approved byyear

als,- declassified1 2 year

< L. J. R o s eAssistant Technical Director0"'61^

S T ,L A v V S , MILETRANSMISSIIN AN Y C . C o r i a n d e r

Technical DirectorT e s t EvaluationIssued as Supplemental Report 2to: Gem ini Program Mission ReportG e m i n i VIIIM3C-G-R-66-1*

by: Gemini VIII Mission Evaluation TeamNational Aeronautics and Space AdministrationManned Spacecraft CenterHouston, Texas

Prepared byMARTIN COMPANY, B A L T I M O R E DIVISION

Baltimore, Maryland 21203Under C O N T R A C T A F 04(695)-394PRI O RI T Y DX-A2

Fo rS P A C E S Y S T E M S DIVISIONA IR FO R CE S Y S T E M S COMMAND

UNIT ED S T A T E S A IR FO R C ELot Angeles, California


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FOREWORDThis 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.

ER 13227-8


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iii

C O N T E N T SPage

Foreword iiSummary vii

I. Introduction 1-1 . System Performance II- l

A. Trajectory Analysis - 1B. Payload Capability 11-39C. Staging 11-39D . Weight Statement 11-41

HI. Propulsion System - 1A. Engine Subsystem - 1B. Propellant Subsystem - 22C. Pressurization Subsystem - 66D . Environmental Control - 77

IV. Flight Control System . IV- 1A. Stage I Flight , . . IV- 1B. Stage Flight IV- 8C. Post- SECO Flight IV- 12

V. Hydraulic 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 (continued)Page

A, Radio Guidance System Performance V-1B. Spacecraft Inertial Guidance System AscentPerformance vi-7

VII. Electrical System Analysis VI-1A. Configuration VI-1B. Countdown and Plight Performance VI-1

VIII. Instrumentation System ; Vni-1A. Airborne Instrumentation . VIII- 1B. Landline Instrumentation Vn-1

IX. Range Safety and Ordnance .................. IX1A. Command Control Receivers IX1B. M I S T R A M IX1C. Ordnance .IX2

X. M alfunc t ion Detection System X- lA. Configuration X- lB. System Performance X- 2

XI. Crew Safety X-1A. Prelaunch Winds Operations X-1B. Slow Malfunction Monitoring X-7

XII. Airframe System XI-1A. Structural Loads -1B. POGO -13


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C O N T E N T S (continued)Page

. AGE and Facilities - 1A. Mechanical AGE - 1B. Electrical AGE - 1C. Master Operations Control Set - 1D . Facilities - 1

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

A. Launch Data Distribution XV- 1B. Film Coverage XV- 6

XVI. Prelaunch and Coun tdown Operations XVI- 1A. Prelaunch XVI- 1B. Launch Countdown XVI- 2

XVII. Configuration Summary XVII- 1A. Launch Vehicle Systems Description . XVn- 1B. Major Components XVII- 3References X V I H - 1Appendix A: Summary of Gemini Launches A- l


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Vll

SUMMARYOn 16 March 1966 Gemini-Titan No. 8 (GT-8) was launched suc-

cessfully and on schedule from Complex 19, Cape Kennedy, Florida.Launch vehicle/spacecraft separation was completed 366 seconds afterliftoff. Spacecraft re-entry was accomplished after completion of 6orbits.The 240-minute countdown was picked up at 0735 EST on 16 Marchand progressed smoothly, with astronaut ingress at approximatelyT-120 minutes. The Atlas-Agena was successfully launched at T-95minutes (1000 EST). The countdown continued normally, and the pro-grammed hold was initiated automatically at T-3 minutes for 5. 9minutes to adjust for proper liftoff time. The countdown resumed atT-3 minutes (1138 EST), and liftoff occurred on schedule at 1141 EST.The spacecraft was inserted into an elliptical orbit with a perigee of86. 9 nautical miles and an apogee of 148. 2 nautical miles, all test ob-jectives for the launch and flight were achieved.Stages I and II engines operated satisfactorily throughout poweredflight. Stage I burning time was 157.9 seconds, with shutdown initiatedby fuel exhaustion. Stage II engine operation was terminated by aguidance command after 182.9 seconds of burning.The flight control system (FCS) maintained satisfactory vehiclestability during Stages I and II flight. The primary FCS was in com-mand throughout the flight. Vehicle rates during Stage I flight did notexceed 2. 5 deg/sec, and the maximum attitude error was 1.89 degrees.The maximum rate and attitude error that occurred during staging didnot exceed 4.2 deg/sec and 1.8 degrees, respectively.Performance of the radio guidance system (RGS) was satisfactory.Pitch and yaw steering signals and SECOdiscrete commands wereproperly executed; ^-^ - ^_^_ _ . - _ -_ _ -_ _ .IGS pitch, yaw and roll performance for the entire flight appearednormal. The dispersions between IGS and primary system attitudeerrors remained within acceptable limits during powered flight.The hydraulic system operated satisfactorily during the 240-minutecountdown and both stages of flight. There were no significant pres -sure perturbations at liftoff or during flight.The electrical system functioned as designed throughout the launchcountdown and flight. Power transfer to vehicle batteries was smooth.


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All channels of the PCM instrumentation system funct ioned satis-factorily throughout the flight. The landline instrumentation systemalso funct i oned satisfactorily prior to and up to l i f tof f . All airborneinstrumentation hold functions monitored in the blockhouse remainedwithin specification throughout the countdown.

The ordnance system umbilical drop weight release, propulsionsystem prevalves, explosive launch nuts and stage separation nutsoperated as designed.The performances of the command control receivers and theM I S T R A M transponder were satisfactory.Malfunct ion detection system (MDS) performance during preflight

checkout and f l ight was satisfactory. There were no switchover com-mands during the flight.The flight environment encountered by GT-8 was within design re-quirements. Flight loads were well within the structural capabilitieso f the launch vehicle. The most .critical loading (which occurred atpre-BECO, af t o f Station 320) reached 103% o f design limit load incompression. _The longitudinal oscillation (POGO) on GT-8 reached a maximumvalue at spacecraft-launch vehicle interface of 0.215 g zero-to-peak

at a frequency of 12.4 cps at LO + 135.4 seconds. Just prior to BECO,another peak of 0 . 295 g zero-to-peak w a s noted f o r 0 . 0 7 second ( 1 o r2 cycles).C r e w safety monitoring, which w a s conducted a t N A S A - M S C , w a sactive during prelaunch and the launch. All guidance monitor param-eters were nominal, and no corrective action was required during theflight.D u r i n g precount operations, propellant loading was delayed approxi-mately tw o hours due to troubleshooting o f heater a n d communication

circuits in the spacecraft. Propellant loading was completed within thescheduled time span and to the specified load and temperature limit.A momentary malfunctionwhich occurred a t T-360 minutes preventedpressurization of the Stage I primary hydraulic system during a rategain test. This did not cause a c o u n t d o w n delay, since the problemcleared up, did not recur and involved components that are used forg r o u n d test only. The T-240 minute countdown was completed onschedule. Acalibration problem with the CP 2650 recorder was.quicklyovercome without delaying the count.All electrical umbilicals disconnected in the planned sequence and

within 0. 825 second. Engine blast and heat damage to the launch standw a s minor.


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IX

GLV- 8 Test Obiectives and ResultsObjective Results

PrimaryP- l Demonstrate satisfactoryboost by the Geminilaunch vehicle system ofa manned Gemini space-craft into the prescribedorbital insertion condi-tions.

P- l Orbit insertion was with-in the predicted toler-ances for V, h and .

P- 2 D emonstrate perform-ance of GLV subsystemsduring powered flight,relative to mission suc-cess and crew safety.

P- 2 All systems performedsatisfactorily throughoutflight. The POGO oscil-lation was 0. 215 g zero-to peak at a frequency of12.4 cps at spacecraft-launch vehicle interface.Secondary

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

S- l Vehicle night was withinthe 3- sigma predictedtrajectory.

S- 2 Demonstrate ability to_ :load propellants^toweight and temperaturelimits imposed by pay-load and vehicle re-quirements.

S- 2 Tanks were loaded withinthe required tolerances"~= =of- weight and- tempera- ture.

S- 3 Demonstrate effective-ness of combined GLVand G A A T V countdownand launch operations,including necessaryground/ range supportsystems to achieve pre-scribed rendezvous mis-sion launch requirements.

S- 3 Atlas- Agena was suc-cessfully launched atT- 95 minutes, followedby the GT-8 l i f toff onschedule. Only a 5. 9minute hold to adjust forproper l i ftoff time wasrequired.


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1-1

I. I N T RO D U C T I O NThis report presents an engineering evaluation of Gemini LaunchVehicle No. 8 ( G L V - 8 ) systems performance during the c ou n td own ,launch a nd powered flight phase o f the Gemini 8 mission.The Gemini-Titan No. 8 ( G T - 8 ) vehicle was launched on schedulef r o m Complex 19, Cape Kennedy, Florida at 1141 hours EST on16 March 1966.Gemini 8 was the eighth mission and the sixth manned flight of theprogram, with astronauts Neil A. Armstrong and David R. Scott aboardthe spacecraft. The mission, which included a rendezvous with theAg e n a Target Vehicle ( A T V ) , wa s completed o n 1 6 March 1966.The GT-,8 vehicle was comprised of the two-stage GLV-8 (similarto G L V - 7 ) and the Gemini 8 spacecraft. T he spacecraft w a s injectedinto an elliptical orbit having a perigee of 86. 9 nautical miles and anapogee of 148.2 nautical miles.Signi f icant events and tests for G L V - 8 at ETR are summarized inFig. 1-1.


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I I - 1

. SYST E M P E R F O R M A N C EA . T R A J E C T O R Y A N A L Y S I S

1. Orbit InsertionGemin i Launch Vehicle No. 8 ( G L V - 8 ) performed as predicted andinserted the G e m i n i 8 spacecraft into earth orbit well with in the allow-able tolerance limits to permit rendezvous with the A g e na target ve-hicle ( A T V) .G T - 8 was steered in the lateral plane during Stage II flight to aset of ephemeris data referenced to the time of insertion (or targeting).

The values of these targeting parameters are given in Table - l; thereare no observed values of these parameters. The targeted and observedinclination angles were 28.868 and 28.92 degrees, respectively. Thetargeted wedge angle at liftoff was - 0. 1245 degree. The observed resi-dual wedge angle at insertion was 0. 0563 degree, which meant t h a t thetotal wedge angle steered was - 0. 0682 degree.TABLE - l

Agena Target Vehicle Ephemeris D a t aG M T L OTR*inVF

60,062.375 sec147, 563.437528.868177 deg- 0.18775463 x 10"525,728.19 fp s rad/ sec

A comparison of the predicted and observed insertion_conditiojis isgiven in Table - 2. In this table and in all succeeding references toa predicted (nominal) trajectory, the data were obtained f rom the G L V - 845- day prelaunch report (Ref. 11), upda ted to reflect the actual space-craft weight (8351 pounds) , guidance constants, T- 0 hour wind and at-mospheric data, and the - 1.34% pitch programmer bias. The observedtrajectory parameters are those derived by the Martin C o m p a n y fromth e Final G E M o d - G 2 pps data. These data have been smoothed an dcorrected for both refraction errors and systematic biases by the Gen-eral Electric Corporation before submittal to the Martin Company.


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I I - 2

TABLE - 2Comparison of Insertion Condi t ions at SECO + 20 Seconds

Al t i tude (naut m i)Inertial velocity (fps)Inertial flight pathangle (deg)

PredictedN o m i n a l86.7342 5 , 7 2 90.001

G E M O D - G86.6672 5 , 7 3 6- 0.020

ObservedM in u sPlanned- 0.067+ 7- 0.021

PreliminaryTolerance 0.394 29.7 Q . 134

2. Derivation of Tralectory UncertaintiesT h e expected ma x i mu m vehicle dispersions and RGS dispersions atBECO and a t S E C O + 20 seconds were obtained f rom Refs. 12 and 13,respectively. A root sum square (RSS) of these dispersions is termedthe preliminary tolerance. After determination of the preliminarytolerance, the total tolerance may be computed by the arithmetic addi-t ion of the preliminary tolerance to the 3- sigma data error of the in-strumentation source being considered. Thus,

/ 2 2Preliminary tolerance = ^[(vehicle dispersions) + (RGS dispersions)Total tolerance = preliminary tolerance + 3- sigma data error.T h e resulting preliminary tolerance is shown in Table II- 3. Becausethe actual insertion conditions were within the preliminary tolerance,th e data error estimates are not needed and, therefore, have been ex-c luded f rom this report.

3. Flight PlanT h e primary objective for G L V - 8 was to place the Ge m in i 8 space-craft into an elliptical earth orbit with an 87-nautical mile perigee*

an d 146- nautical mile apogee. * Having achieved orbital insertion at25, 730 fps, ** the spacecraft t h e n separates f rom Stage II (adding 10 fpsto spacecraft velocity in the process) and coasts to the desired apogee.The following flight plan was employed to attain the desired conditions.A vertical rise is planned for the first 23.04 seconds following l iftoff,during which time a programmed roll rate of 1.25 deg/ sec is initiatedto roll the vehicle f rom a pad orientation of 84. 933 degrees to the flightazimuth of 99. 9 degrees.

Relative to Complex 19.**Does not include the separation velocity imparted by the spacecraft.


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At this time, an open- loop pitch program is begun (via a three- steprate command) which terminates at 162. 56 seconds. The nominal com-manded pitch rates and their times,of application are shown in Table II- 4.Guidance commands from the radio guidance system (RGS) areinitiated at LO + 168. 35 seconds and continue until two seconds priorto SECO, however, velocity cutoff computations continue to SECO.Between SECO and SECO + 20 seconds, the engine shutdown impulsecontinues to add velocity to the vehicle (approximately 81 fps), and thespacecraft is separated from the sustainer after SECO + 20 seconds.

TABLE II -4Planned GLV Pitch Program

ProgramStep 1Step 2Step 3

Rate(deg/ sec)- 0. 709- 0.516- 0.235

Time from Liftoff(sec)23. 04 to 88. 3288.32 to 119.04

119.04 to 162.56 .A comparison of the planned and actual sequences of events is con-

tained in Table II- 5, and a profile of the G TA- 8 flight superimposedon the range planningmap appears in Fig. II- 1.4. Trajectory Results

Analysis of the range data and Mod -G radar data indicates thatthe performance of G LV- 8 was normal and the vehicle flew close tothe prescribed ascent trajectory throughout Stages I and II.Table II- 6 shows a reconstruction of the BECO condition by con-sidering the actual engine data, weather conditions, propellant loading,

engine misalignments, wind and guidance errors. This table is com-prised of those items which can be measured and those which can onlybe estimated due to lack of suitable instrumentation. The primaryfactors contributing to the pitch and yaw plane trajectory dispersionsat BECO are listed and the effect of each is summarized.A Y_ coordinate displacement at BECO results in an approximate

amplification of five on the YF coordinate displacement at SECO + 20seconds. The differences shown in the apparent and measured incre-ments of Table II- 6 are well within the allowable tolerance limits pre-sented in Table II- 3.


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- 5

S/ c separation (526 ,075) j

Fig. Il- l. GT- 8 Boost Plight Path Profile


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- 6

TABLE - 5G T - 8 Flight Events Summary

Measurement0800/0801FCB-10210403560357210101694421442244230734073407320732073207280732073207350741035603570032050201690855073207400755/ 075607390777051905220521079908550699

EventPower TransferMOCS T-087FS (T-0)Stage I S/ A- 1 MDTCPS MakeStage I S/ A- 2 MDTCPS MakeTCPS S/A- 1 & S/ A- 2Launch NutsFirst Motion 'Shutdown Lockout (backup)Liftoff 'Start Roll ProgramEnd Roll ProgramStart Pitch Program No. 1

- Stop Pitch Program No. 1Start Pitch Program No. 2PCS Gain Change No. 1Stop Pitch Program No. 2Start Pitch Program No. 3Staging Enable (TARS discrete)IPS Staging Arm TimerStage 1S/ A- 1 MDTCPSBreakStage I S/ A- 2 MDTCPS Break87FS,/ 91FSX (BECO)Start PC- RiseStage .SeparationStage II MDFJPS MakeStop Pitch Program No. 3R G S EnableFirst Guidance CommandStage E Shutdown EnableGuidance SECO91FS2Shutdown Valve RelayShutdown SquibASCOStage MDFJPS BreakSpacecraft Separation

G M T(hr- min- sec)1639:33.41640.59.037

59. 0881641.00.030

:00. 021:00. 098:02. 206:02. 280:02. 29002. 389.10. 87

, - 22.8625. 43

1642 30. 63:30. 63.47. 151643 01.2601.2626.8026.96:36. 969 36. 98037.00437. 650

37. 68837. 661

.44. 1 144.0450.79

1646:18.68.39. 90539. 925

.39.96:39.9539.97:40. 065

1647 08.048

Time from Liftoff (sec)Actual- 89.0- 3. 352- 3.301- 2. 359- 2. 368- 2.291- 0. 183- 0. 109- 0. 099

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20.4723.0488.2488.24

104. 76118.87118.87144.41144. 57154. 580154.591154.615155.261155.299155.272161.72161.65168.40316.293371516337.536337. 57337. 56337.58337. 676365. 659

Planned- 89

- 3.46- 3.37- 2. 27- 2.27- 2.20- 0.20- 0.10- 0.10

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104. 96119.04119.04144. 64145.00153. 79153.79153.85154.50154.58154.75162.56162.56168.35317.44336.73336. 75336.77336.77336. 78337.07356. 75


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II-8Table II-7 presents the trajectory parameters computed from theGE Mod III-G, MISTRAM I and MISTRAM I and II data. At BECO andinsertion, the two data sources yielded comparable results with theexception of the crossrange position (YF) and the crossrange velocity(Y_,) parameters as both flight events,

The actual, as well as the predicted, nom inal trajectory is pre-sented in graphical form in Figs. II- 2 through 11- 24. On these graphs,the nominal trajectory is that documented in Ref. 11, updated to reflectthe actual spacecraft weight (8351 pounds), guidance constants, T- 0hour wind and atmospheric data, and the - 1. 34%pitch programmer bias.The observed flight data were obtained from the Mod III-G 2- pps data,smoothed and corrected for refraction errors an d systematic biases.A list of the primary tracking sources with the tra jectory time inter-val covered by each is contained in Table II- 8.

5. G eodetic and Weather ParametersSignificant geodetic and weather parameters are shown in Table II- 9.The atmospheric pressure an d temperatu re variation with altitude isdepicted in Fig. 11- 25,"the pressure arid temperature were essentiallystandard. F igure 11- 26 presents the altitude history of the magnitudeand direction of the wind. "At low altitudes the winds were light, in-creasing to a peak of 125 knots at 37, 500 feet. The wind was nearly aside wind up to approximately 7,000 feet then shifted to essentially atail wind.

6. Look AnglesThe maximum look angle in pitch (LAP) occurred at LO + 336 sec-onds, when it attained a value of 20. 7 degrees. This maximum valuewas within the boundary existing at that time, as shown in Fig. 11-27.,T,he corresponding look angle in yaw (LAY)was also within the estab-lished limitation, as shown in Fig. 11-28. The maximum value of LAYwas 4.8 degrees, which occurred 159 seconds after liftoff.

TABLE - 8Data Available for Trajectory Analysis

SourceAFETR

GE

NASA- MSC

TypeMISTRAM posi-tion, velocityand accelerationMod III-G radar

"position, velocitySpacecraft IG Sascent param-eters

StationValkaria IEleuthera II

Cape Kennedy

Flight Coverage(sec from range 0)65 to 374. 8132. 4 to 374.8

LO to + 424

LO to + 380


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- 9

TABLE II-7Comparison of GT- 8 Predicted and Observed Performanceat BECO and at SECO + 20 secon ds

(sec)veloc i t y (fps )(f t )f l ight p ath angle (deg)r ange ( n a u t mi)

radius (f t )position X (f t)

YF (ft)pos i t ion, Zp (ft)

veloc i t y, (fps)YF (fps)

, Z_ (fps)Vy (fps)

steering veloc i t y, Vy (fps )seconds

LO sec)veloc i t y (fps )(f t )f l ight p ath angle (deg)r ange (n au t mi)

radius (f t )position, X_ (f t )position, Y_ (ft)

position, Zf (f t )ve loc i t y , (fps )

YF (fps )velocity, Z p (fps )

ve loc i t y , Vy (fps )steering velocity, Vy (fps)

Predi c t edN o m i n a lTraj ec t ory(R ef . 11)*

153. 8479970. 8207.556.19. 17350. 93821,117,565.31 2,7 7 0.2695.205,211.8128.25. 13151.-- --- -

356.48625 ,729 . 1527,005.0. 0005538. 33021,438 ,572 .3 .340 ,309 .77,938 .265 ,329 .24,045.942.0-3 ,805 .3. 26- 1 . 2 6

Tracki ng Fac i li t yGE Mod I I I - G(2 pps)

154.6159917.4209 ,005.19 .3450.9321,119 ,011.312 ,690 .37 64 .206,660.8067 .56.83161.213.-- -

357.53625 ,736526.600.-0 .0200538 .6821,438 ,152 .3 ,342 ,304.84, 05T.264.585 .24,049 .960.2-3,818.-16 .0

MIS T R AM I

154.6159906.7209,081.19.3450.93121.119 ,085 .312 ,7094278.2 0 6 , 7 3 78056.69 .831 59.......357. 53625,735526,951-0 0110538. 5921,438 ,488 .3,342,158.89,811.264,937.24 , 04 7996.9-3 ,814....-- -

MIS T R AM Iand II

154.6159910.209 ,06619.3450.93421,119 ,070.31 2,7 20.4 286.206.720.8057.703158....-- -

357. 53625,7 37526,248-0 . 0448538.712 1 , 4 3 7 , 7 8 73,342 ,217.89 ,878 .264,218 .24 .04 6997.-3 ,829 .-- -...

BET

154.6159918 .2 0 9 , 0 7 419.3550.93321, 119 ,078 .312,7164282.206,728806770.3163.-- --- -

357,53625,7365 2 6 , 4 7 2-0 .0330538 .8821,438 ,0083 , 3 4 2 , 2 0 089 ,8182 6 4 , 4 4 624.046.997.- 3 8 2 4 .-- -

weight (8351 pounds), gui dance cons t ant s , T -0 houra t mospher i c da t a , and t he -1 . 34% p itch programmer bias.

ER 13227-8


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Predicted n o m i n a l wi nd ru n 80 -GT-6 ( f in al )*G. E. M o d III -G f in al f l ig h t data*

:*Include si R a w i n s o n d e balloon d ata! Cape K en n ed yi 1011 EST, 16 March 1966

P r e d i c t e d B E C O '_/(15 3 .847 sec) ft/ .. . :BECO(154.615 sec) j

10 20 30 40 50 60 70 80 90Time f rom Li f t o f f (sec)

100 110 120 130 140 150 160 170 180

Fig. II-2. Inertial Velocity Versus Time: Stage I Flight

[ A L


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** / :.." I"** ; /.* : : I ! . - ! ( : ' ! ! \ 1 ! 'I j p j 1 j ' '[ Predicted BECOL/.jj(153.847 sec) !: ' . - j M . - : ( ; 1 ; . | : : i i . 1 .:.j ; 1 '

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

ER 13227-8


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vg

Predicted n o m i n a l w i n d run 8 0 -G TA - 8 ( f i na l ) *G . E . Mod I I I -G f i na l night data*

R a w i n s o n d e balloon dataC a p e K e n n e d y1011 EST, 16 March 1966;!: .I".

BECO' (154.615 sec) i

10 40 50 60 70 80 90Time from Li f t o f f (sec)

100 110 120 130 14 0 150 160

Fig. 1110 Axial Force Versus Time: Stage I Flight

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90

~ 70 50

,.

40

30

3V< 20

10

/9 - / lauaeJb. IDUIIHL NTIALI J - 1 9

*Includes

:

Predicted nominal wind run 80-GTA-8 (final)*G .E . Mod III-G final night data*

e balloon dataedy16 March 1966

-p p ;~r

X*f '

.. * * '

.' 1

.

11I 1

V

/A1/

JL,1P r e d i c t e d B E C O( 1 5 3 . 8 4 7 sec)

'.*

.

Y - : - ,

B E CO( 154. 615

-. -

Bl )

//. ,

.

0

!

10 20 30 40 50 60 70 80 90 100 110 120 130 140Time from Liftoff (sec)

Fig. 1111 Aerodynamic Hea ting Indicator Versus Time:

150 16

Stage I Fligh

IDC ER 13227-8

J_


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- 20

3 10

-

-40

_....

10

P r e d i c t e d n o m i n a l w in d ru n 80- G T A- 8 ( f ina l) *G . E . M o d III -G f ina l n ight data*

[

* Inc ludesR a w i n s o n d e ba l loon da taCape Kennedy1011 EST, 16 M arch 1966P redic ted B E CO= ( 153 . 847 sec)

..- r ,...** * * % * .

B E CO( 154. 615 sec)

70 80 90 Time from Liftoff (sec)

120 130 140 150 160 170

Fig. 1112. Stage I Angle of Attack History

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- 21

4'tl'l20

LO Predicted BECO(153.847 sec)

. *%. ;

Predicted nominal wind run 80-GTA-8 (final)*G.E. Mod III-G final night data*

-40 ^IncludesRawinsonde balloon dataCape Kennedy1011 EST, 16 March 1966

-60 . . I , 51 80 90 100Time from Liftoff (sec)

Fig. 11-13. Stage I Angle of Sideslip History

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11-2228

26

5|gd1

Predicted SECO + 20 (356.486 sec)

SECO + 20 (357. 536 ) . iMii:IncludesR a w i n a o n d e balloonCape Kennedy1011 EST, 16 March

140 160 180 200 220 240 260 280 300 320Time from Liftoff (sec)

340 360 38 0 400 4? 0 440 460

Fig. 1114. Resultant Inertial Velocity (V ) Versus Time: Stage II Flight

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*!2

20h \

N

r

:

-

12

'/

Predicted nominal wind run 80-G TA-8 (final)*G.E. Mod III-G final flight data*

:

' 1 '

- t L ; ! , ;

:

I *1 (R a 10'

Rawinsonde balloon dataCape Kennedy1011 EST, 16 Marc h 1B66

r j_

SECO + 20 (357. 536 sec)

; !Pred icted SECO + 20 (356.486 sec)

I ! 211?0 140 160 180 200 220 240 260 280 300 320 340 360 380 400 420 44 0 460Time from Li f t o f f (sec)

Fig. 11-15- Inertial Flight Path Angle (y.) Versus T i m e: Stage II Flight 1

I I - 23-z

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Predicted SE CO + 20 ( 356 . 486 sec)"1P redic ted nomina l wind run 80 - G T A- 8 ( f ina l )*G . E . M o d I H - G f in a l night data*

!

* Inc ludeaR a w i n s o n d e balloon dataCape Kennedy SE CO + 20 (357. 536 sec)1011 EST, 16 March 1966

l . ' . i f - ; t ; - i : m " liajt

160

1 '

:120 140 180 200 22 0 240 260 280

Time f rom Li f t o f f (sec)300 320 340 360 380

Fig. II- 16. Altitude Versus Time: Stage II Flight

CONFIDENTIALER 13227-8


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4.0

3.6

~ 3.2IB

Xi lTJU

2. 8

2. 4

2. 0

1.6

Predicted S ECO + 2 0 (3 5 6 .486 sec)

S ECO + 2 0 (3 5 7 .5 3 6 sec)

fl*In eludesR aw in so n d e ballo o n d ataC a p e K e n n e d y10 11 ES T, 16 M arch 19 66

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 A - 8 ( f i n a l ) *G . E . M od I I I - G f i n a l night data*

320 340 360 380 400 420 440Time from Lif to ff ( sec)

Pig. 11-17. Downrange Position Coordinate (X.,) Versus Time: Stage Flight

11-25

* ER 13227-8


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160,

140

120

2

4j 80 ~i_L_

20

-20:

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 A - 8 ( f i n a l )*G . E . M o d I I I - G f i n a l night data*

J * I n c l u d e sI R a w i n s o n d e b a l lo o n d a t aI C a p e K e n n e d y1011 EST, 16 M arch 1966h i S E C O + 2 0 ( 3 5 7. 5 36 sec)

; :

IrfM f^; :

-


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.. ' /TIAL

52 0

48 0

440~400

360

\I " : :; 1 I - ' : ! ' '

;: I320

280

24 0

200

160

120

Predicted n o min al w in d ru n 8 0 - G T A - 8 (final)*G . E . Mod III -G f ina l night data*

.R a w i n s o n d e ba l loon dataCape K en n ed y1011 EST, 16 March 1966

SE CO + 20 (357. 536 sec)

Predicted S ECO + 2 0 ( 356 . 486 sec)

.'

200 220 240 260 28 0 300 320 340 360 380 400 420 440 460 480Time f rom Liftoff (sec)140 160 180

. Fig. 11-19- Vertical Position Coordinate (ZF) V ersus Time: Stage Plight

F l13227-8


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11-28 - IDENTIC

: ! ,| ! Predicted n o min al w in d ru n 8 0 - G T A - 8 ( f i n a l) *G . E . M od I I I - G f i n a l night data*HT

S ECO + 2 0 (3 5 7 . 5 3 6 sec)

J . - L - -iH

ffli : : i : jPredicted S E C O + 20 (356.486 sec) imcludesR a w i n s o n d e balloon dataC a p e K e n n e d y1011 EST, 16 M ar ch 19 66 - - + - 1

- -- --

160 180 200; - ;260 280 300 320

Time f ro m L i f to ff ( sec)440

Wg. H-20. Stage II ngle of ttack History

CONFIDENTIALER 13227-8


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- 29

S ECO + 2 0 (3 5 7 .53 6 sec)

Predicted S ECO + 20 (3 5 6 .486 sec)

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 A - 8 ( f i n a l ) *G . E . M od I I I - G f i n a l night data*

- 16 *In clu d esR aw in so n d e 'bal lo o n d ataC a p e K e n n e d y1011 EST, 16 March 1966220 240 260 280

Time f ro m L i f to ff ( sec)320 340 360

Fig. 11-21. Angle of Sideslip Versus Time: Stage II Flight

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IT- 30- I 11-30-

1UU

50

~_iX5 to

*

214.3

1.0

0.5

aII- 1 ? .

h4) -0.5

-1.0

25. 85

25. 80p

1tf0

X 25.75!-x ....-2V> ah )

25. 70

25.65 -3

334 336

Cross-ran ge velocity

Cross rangeilnertial velocity1

Inertial flight path angle

,- -i"V\ .Ground range

SECO + 20sec\,S E C O '338 - -340 344 346 348

Time from Liftoff (sec)350 352 354 356 358 360

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

CONFtDENftAkER 13227-8


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^ 11-31

1000

P redic ted SE CO + 20 ( 356 . 486 sec )P r e d i c t e d n o m i n a l w in d ru n 80- GT A- 8 ( f ina l ) *G . E . M od I I I -G f ina l n ight da ta*

.* Inc ludesR a w i n s o n d e ba l loon da taCape K e n n e d y, 1011 E ST , 16 M a r c h 1 9 6 6 :

S E C O + 20 (357.53 6 sec)

- f !

i_

120 140 160 180 200Time from Liftoff (sec)

220 240 260 280 300 320 340 360

Fig. 11-23. Cross-Bange Velocity (Y.-J Versus Time


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'

800

11-32 ^CONFIDENTIAL

:

700

600

.-i__iU

500

f r i '- b400

75 300

-T-j-1 1..; . ......

~[ ;200

i U100

.j

;

E.

: - -

- - 1.. .

~ i . ] ' -

H is

- -. -

- : . . ! - .*

.~r

: :

Z E

:;r:::

... :....!.-. ',.;:

-

11

-~ ~

"""""" *i

:i:t . ; . . . .

l ; l ! l l ' ! l !I H M r i fr ' ,./

7Tj~ 1- 1

| . : : ' l : ;-

:; 1 > v'i ': ,h_. 1: :(:"I ! ' ! - ' : ' - ! ' . - ' . -i - - i : : : : |::::f:::' ' ' ' . ' . . . : - 1-\: : : i . . -P r e d i c t e d n o m i n a l w in d run 80 - GT A- 8 ( fina l ) *G. E . M od I I I - G f ina l f l ight data*

I"0*

* Inc ludesR a w i n s o n d e ba l loon da taC a p e K e n n e d y1 0 1 1 EST, 16 March 1966I/"

ss. . i :

J~r ^ - . ~ "~

f~::.J_.

-i ;: ) | "*^-o ," - - - ,

:-]_

.

: '[' -': i

|' " " - ' " :

"

.

: : ( ; - ; . .

: ,! .: :

\ -.J : : '

i

'

"|. ' . 4 ~

~1~' / : ] . 4~] 11.; { : ! : *~ !f~

: !.' .-..: : 'j -

f" '.;; j: :;

:-|- : . i : :' .

;. !: : ; ;.;^: r - i ; . r

~"

fa-- ". : . :

~^-~' - - : ' - 1

:. :i : . .:: -:|...

.. .

--

_

4 J H -U 'j; ':::::.];::.

1, .' : . '' ' ! : ' : ::

~ . ' \ ' -~: ( : . .a ii rtrgr:': ':. i: ':;::!: '

| Predicted SE CO + 20 ( 356 . 486sec){ ' - I " - 1 : "iil - - : : [ .

*( i| * -

tabhL

!^~**

|

^ ^

, . - _ L ilX^

-'

? *sS b ^ ~ -

1 !

. - : .

-::

.-:; ;:

-

:.. 1 .'

-. . .

.-||-: :: ; '. : -::t- : . : . [ . -.: -.... .:-;: - -\iHEii- - J - f H - - :1. .. . -}-'...1_. ' " I : : . ' : ; ! .::;;{ .ijrfc: : . 'I;V1 1-. : 1 . !^ r , f :.-"'}Si* 4-

-100 fr

i- 2 0 0 20 40 b U 80 100 120 140 160 180 200

T ime f rom Liftoff (sec)

SE CO+ 20 ( 357. 536 )

26 0 280 300 320 340 360

Fig. 11-24. Yav Steering Velocity (Vy) Versus Time

IDENTIE R 13227-8Q


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-

G eographic and Weather Conditions at LaunchLocation

SiteSite coordinates:

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

Complex 19

28.507 N80.554 W84. 933 true azimuth

WeatherAmbient pressure (psi)Ambient temperature (P)Dew point (F)Relative humidity (%)Surface wind:

Speed (fps)D irection (deg)

Winds aloft (max):Altitude (ft)Speed (fps)D irection (deg)

Cloud cover

14.76705969

30350

37,500211268 true azimuth0. 3 strato- 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 toellipsoid and J_ X- axisForms a right- handed orthogonalsystemFischer

LaunchInitial flight azimuth (deg)Roll program (deg)

99. 9 true azimuth15. 0 ccw

S


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-34 COMF1

100

*IncludesR a w i n s o n d e balloon dataCape K e n n e d y1011 EST, 16 March 1966

0 20 40 60 80 100 120 140 320 360 320 280 240 200 160 120 80W in d Speed (kn) WndAzimuth(degfromnorth)

Fig. 11-25. Wind Speed and Azimuth Versus Altitude


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;U-a

'- 1.->I

ER18


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t

FDN


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- 38 CONFIDENTIAL

7. Maximum Dynamic PressureThe maximum dynamic pressure for the GTA-8 trajectory was lessthan design limits. Table - 10 compares the predicted and observedconditions associated with the maximum dynamic pressure. The pre-dominantly tail wind environment for this flight in itself reduces themaximum dynamic pressure. A predicted t rajecto ry computation fora no- wind condition showed that the maximum dynamic pressure wouldbe 743. 4 psi, and the predicted trajectory with T- 0 winds, from Table11-10, shows a value of 681. 3 psi, verifying the effect of a tail wind.

TABLE - 10Trajectory Parameters at Maximum Dynamic Pressure

Dynamic pressure (psf)Time from liftoff (sec)Mach numberAltitude (ft)Relative flight path angle (deg)Relative wind velocity ( fps)Wind velocity (fps)Wind azimuth (deg from north)Angle of attack (deg)Angle of sideslip (deg)

Predicted*(nominal)681.379.621.7345,00047.521665.1352660.35

- 0.09

Observed**678.380. 121.7445,30647.4516731362670.65

- 0.35

*Ref. 11, updated (see footnote to Table II- 7)**Mod III-G 2 pps radar data


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II-39

8. Angles of Attack and SideslipPredicted 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 andatmospheric information obtained from the 1011EST Rawinsonde sound-ing. Observed angles of attack and sideslip were derived using the ModIII-G position and velocity information, IGSattitude data and the afore-mentioned weather data.

B. PAYLOAD CAPABILITYPropellants remaining onboard after Stage II low level sensor un-cover indicated that a burning time margin (BTM)of 1. 327 secondsexisted to a command shutdown. The total propellant weight marginwas 437pounds, and the corresponding GLVpayload capability was8826 pounds. These values and the predicted nominal and minimumvalues appear in Fig. 11-29. The predicted capability curves were ob-tained from the real-time propellant temperature monitoring digitalprogram (Run 14) adjusted to reflect the pre-liftoff temperature changesand the actual Agena ephemeris data used in the guidance equations atlaunch. The predicted propellant weight and burning time margins arebased on the difference between these curves and the 8351-pound

spacecraft weight.The last payload prediction indicated that the minimum payloadcapability was 215 pounds less than the spacecraft weight, and thenominal payload capability was 393 pounds greater than the spacecraftweight at the predicted launch time. The actual (postflight reconstructed)GLV capability was 475 pounds greater than the spacecraft weight.

C. STAGINGThe staging sequence was normal and physical stage separation oc-curred as planned. The time interval from staging signal (87FS9/91FS1)

to start of Stage II engine chamber pressure (P ) rise was 0. 646 second.C3This compares favorably with the nominal expected time of 0. 70 - 0. 08second. Stage separation occurred 0. 029 second following start ofP rise.C3.

.


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II-40

A"a

M

c d > gI's

G T - 8 Flight Test Value

200- 3 0 0- 4 0 0 1400300200100

- 100- 200- 300

- 40021

- I

i b i : 7 l ; . ; : ; . - Nominal -r t f

ight = 8351 lb'

: < - I > 1

' : -: . ! : . : 1 ' ' Minimum I ff

: : !" - L

:::: -! ^B p i

_.- - ..

: : :, .

G T - 8 F l i g h t Test V a l u e |Nomnal _~ , - ^ = T ; ,] M in im u m

Phase pane - H - i : : : : ; i20 40 60 80 100 120

Time i n L a u n c h W i ndow (min)Fig. 11-29. Payload Capability

140


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- 41

D. WEIGHT STATEMENTTable 11- 11 shows the G T- 8 weight history from laun ch to orbitalinsertion.The postflight weight report (Ref. 12) provides the background datafor this summary. The report includes a list of dry weight emptychanges at ETR and shows a derivation of weight empty from the actualvehicle weighing. Other items covered include the derivation of burn-out, BECO, SECO and shutdown weights; weight comparisons with theBLH data; and the center of gravity travel envelope as a function ofburn time for the horizontal, vertical and lateral planes.

TABLE II- 11G T - 8 Weight Summary

Loaded weightStart and grain lossesTrajectory LO weightPropellant consumedto BECOCoolant waterWeight at BECOShutdown propellantWeight Stage I burno utStage II engine startStage II LOPropellant consumedto SECOAblative, covers an dcoolant waterStage II at SECOShutdown propellantWeight at SECO+ 20 seconds

Weight (Ib)Step I

271,366- 3,688(1)267,678256,384

011,294

35310,941(3)10,941

Step II65,642

65,642110

65,631

65,631(3)192

65,43959,422

42

5,975136

5,839

Step III8.351(2)

8,351

08,351

8,351

8,351

8,351

8,351

Stage Total345,359

341,671

85,276

84,923

73,790

14,326(4)

14. 190(4)

NOTES:(1) Event: TCPS + 2 seconds (launch bolts blown)(2) Information from N ASA- H ouston (3/ 15/66)(3) Includes outage: 252- lb Stage I, 261- lb Stage II(4) Includes 437 Ib of burning margin


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ENTIAL- . P R O P U L S I O N SYSTEM

A. E N G I N E SUBSYSTEMThe Stages I and engines operated satisfactorily t h r o u g h o u t theflight, an d all l a u n c h objectives were met. Stage I burn ing time was157. 918 seconds, and shutdown was initiated by fue l exhaustion. StageI I operation was terminated by a guidance command at 91 PS. + 182. 921seconds. Several anomalies occurred during the flight, n on e of w h i c h adverselyaffec ted engine performance. These were:

(1) The Stage I oxidizer pressurant pressure switch (OPPS)cycled o n c e prior to f i na l actuation at + 1. 674 seconds.Approximately 7. 5 milliseconds after initial actuation theswitch deactuated for 7. 5 milliseconds before f i na l "make. "Pressure fluctuations were not observed in the oxidizerpressurant orifice inlet pressure (POPOI), although thelo w sampling rate may have obscured a perturbation. It isassumed that the OPPS chattering on G L V - 8 was a randomevent. Chattering of OPPS c a n n o t i n f lu e n c e an engine shut-d o w n prior t o t he interrogation time o f + 2 . 2 seconds.(2) The Stage II thrust chamber pressure start transient indicatedan abnormally slow rise rate. The most probable cause wasa temporary obstruction in the sensing line or transducercavity.(3) Stage I I engine performance began to decay an d exhibited anunusual time- dependent characteristic after 91FS. + 90 seconds.

The u n u s u a l operation was predominant in oxidizer flow rate,mixture ratio and thrust. Examination of flight data indicatesabnormal oxidizer p u m p operation.(4) Five disturbances were no ted in accelerometer data afterSECO and before spacecraft separation. The disturbanceswere not characteristic of the post- SECO disturbances ob-served on previous GLV and Titan II flights. A correlationof the disturbances wi th thrust chamber pressures c a n n o t beestablished. A post- SECO disturbance, characteristic ofprevious flights, was observed after spacecraft separation.

The only significant configuration change a f fec t ing performance wasthe incorporation of the Stage Gemini Stability Improvement Program( G E M S I P ) th rust chamber injector. Analysis of the flight data indicated


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Ill- 2

satisfactory performance of this injector. Although several anomaliesoccurred during Stage operation, n o n e appeared related to injectorperformance.1. Stage I Engine ( Y L R 8 7 - A J - 7 . S / N 1008)

a. C onf ig u r a t i on and special proceduresT h e G LV- 8 Stage I engine di f fe red from that o f G L V - 7 by the retro-fi t replacement of the rigid tube between th e lubricating o il heat ex-changer a n d t h e fue l p ump with a flexible line in order to minimize ex-cessive vibration and tube failures (Aerojet ECP- 076).T h e th rust- chamber valve through- bolts were torqued to only 100i n c h - p o u n d s a t M a rtin- Baltimore, then retorqued t o t h e final 300 to 320i n c h - p o u n d s at ETR . This procedure was invoked to preclude a possiblestress corrosion problem.All propulsion system breaks of integrity were reviewed to verifyt h a t procedures were adequate t o ensure that closures h a d n o t inad-vertently been left in t h e engine during reassembly.T h e pressure sequen ce valve (PSV) drain line engine attachm entconf igura t ion was modified to prevent a recurrence of the prematureseparation of the drain lines from th e vehicle w h i c h occurred during th e

G T - 6 s h u t d o w n . O n t h e G T - 6 l aunch attempt, th e resultant fire i n thef l ame pit area created a hazard and delayed astronaut egress. The slip-fi t polyethylene tube to the engine drain line as previously used wasreplaced with a clamped configuration incorporat ing a weak link V - n o t c hin th e polyethylene tube.b. Start transientT h e S / A 1 a n d S / A 2 thrust chamber pressure start transients werenormal and are presented in Figs. - l and III-2. The ignition spikesi n d i c a t e d 88% (687 psia) and 81% (634 psia) o f rated thrust fo r S /A 1

an d S/A 2, respectively, which are above the engine model specificationallowable (75%). However, the GLVengine chamber pressure instru-m e n t a t i o n h a s characteristically shown u n d a m p e d oscillations which o b -scure the true transient performance and prevent accurate determi-nation of the ignition spikes.Significant start events are presented in Table - 1.


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1000

800

/ " FIDENTIAL

M D TC PS (M eas 0 3 5 6)|

+ 2 . 5 +3.0

Fig. III-l. S/A1 Start TransientTime from 87FS. (sec)

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1000 - 4 - TWPFNTIA

800

-I\600

400

200 -

*fyJVWt^

In u - - ^ .- . ; .' .! ;. ,, ,- . - - ! t

MDTCPS (Meas 0357) -:

. , 1 1 , I I H f ' )

+ 2 . 0 + 3 . 0Time f rom 87FS (sec)

an 11Fig. IH-2. S/A 2 Start Transient

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CONFIDENT - 5

T A B L E III-1Stage I Engine Start Parameters

ParameterF S X to initial P rise (sec)P ignition spike (psia)P step- - approximate (psia)P overshoot (psia)

S / A 10 . 7 7 3687455N o n e

S / A 20 . 7 6 6634440N o n e

c. Steady- state performan ceStage I engine flight performance agreed closely wi th th e preflightprediction. Flight integrated average performance parameters werewith in 0. 8% of th e preflight predictions.From measured flight data, engine performance was calculatedwith th e Martin- Baltimore PRESTO computer program, using th eStage I thrust coef f ic ien t relationship as modif ied by Martin- Baltimore.T h e modi f i ca t ion increased thrust an d specific impulse by approximately3400 p o u n d s a n d 2 .0 seconds, respectively, above t h e values calculated

u s i n g t h e Aerojet thrust coeff ic ient relationship. T h e Martin- Baltimoremodi f i ed thrust coeff ic ient w a s also used fo r preflight predictions.T h e Stage I engine average flight performance, integrated f romliftoff to 87FS2 is compared with the preflight prediction in Table - 2.

T A B L E - 2Predicted an d Flight Performance Comparison- - Stage I Engine

ParameterThrust, engine (Ib)Specific impulse,engine (sec)M ixt u r e ratio, engineOxidizer flow rate,overboard (Ib/ sec)Fuel flow rate, over-board (Ib/ sec)

PreflightPredicted Average*462, 508278.141.94431097.78565.11

FlightAverage*4 6 1 , 2 3 3278.511.92901090.33565.74

D i f f e r en c e(%)- 0.28+0 .13- 0.79- 0.68+0.11

*Martin- Baltimore modified thrust coefficient relationship used.


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Ill-6

T h e eng ine pe r fo rm an ce ca lcu la t ed th ro ugho u t the Stage I f l ight ispresen ted in F ig. - 3. The prefl ight prediction is shown for compari-son .O n G L V - 8 , t h e t h r u s t c h am b e r pressure t r a n s d u c e r s w e r e w r a p p e dwith t w i c e th e n o r m a l a m o u n t o f t he rmal i n su l a t i on in order to m i n i m i z eP i n s tr u m e n t a t i o n d r i ft . R e c o n s t r u c t e d fli gh t d a ta sh o w e d t h a t t h e

d r i f t was de l ayed f rom the p rev ious average t ime o f approx imate ly 87FS ,+75 s e c o n d s to a p p r o x i m a t e l y 87FS1 + 90 seconds . A l t ho u gh P dr i f t fo rth e Stage I engine subassemblies r e a c h e d - 2 . 8 % o n this f l ight as com-p a r e d to a prev ious average of - 1. 5%, it is r e c o m m e n d e d th a t the add i -t io n a l w r a p be c o n t i n u e d fo r the r emain ing f l i gh t s .T h e pre f li gh t p r ed ic t ion fo r G L V- 8 was the first to use the f ina la c c e p t a n c e test t i m e - d e p e n d e n t biases directly rather t ha n to app ly thea c c e p t a n c e - t o - f l i g h t f ami ly co r r ec t i on (de r ived f rom T i t an I I da ta ) .G L V f l i gh t exper i ence has shown better c o r r e la t io n w i t h a c c e p t a n c ebiases. Figure 1- 3 sh o w s th e close t im e - d e p e n d e n t a gr e e m e n t o f t h ef l ight results wi th th e p re f l igh t p r ed ic t i on .

Stage I eng ine f l i gh t pe r fo rmance ca l cu l a t ed a t the 87FS . + 5 5 - s e c o n dt i m e slice and c o r r e c t e d to s t andard in l e t cond i t i ons is sho w n in Table 1- 3. T h i s is c o m p a r e d w i t h the a c c e p t a n c e test and the pred ic t ed f l i gh tp e r f o r m a n c e , at s tandard in le t condi t ions , and the n o m i n a l t im e as usedin th e pre f l i gh t p r ed ic t i on . T h e pred ic t ed f l i gh t pe r fo rmance a t s t andardc o n d i t i o n s was ob ta ined by m odi f y ing t h e n o m i n a l a c c e p t a n c e test da tafor a 4 8 5 0 - p o u n d a c c e p t a n c e - t o - fl ig h t t h r u s t g r o w th o b t a in e d f r o m a n a l y-sis of prev ious T i t an II and GLV f l ights. The 5 5 - s e c o n d t i m e slice p e r -f o r m a n c e c o r r e c t e d to s t andard in l e t cond i t i ons is not r e p r e s e n t a t i v eof t h e total a c c e p t a n c e o r fl ig h t p e r fo r m a n c e ; h o w e v e r , i t is t h e n o m -ina l t ime used fo r quo t ing r a t ed pe r fo rm an ce . T he da t a , t h e r e fo re ,a r e i nd ica t ive o f t he co r r ec t ed pe r fo rm an ce a t 55 secon ds on ly becauseof variations in the t i m e - d e p e n d e n t biases an d eng ine in l e t cond i t i onst h rou ghou t f l i gh t .d . Shutdown t r a n s i e n t

Stage I engine sh u tdow n was in i t iated by fue l exhaus t ion . Figures - 4 an d 1- 5 sh ow t h e S / A 1 a n d S / A 2 c h a m b e r pressure decays.All engine parameters w e r e n o r m a l for a fue l exhaus t ion shu tdown .T h e eng ine th rus t at staging was approx imate ly 57, 000 pounds. S ign i f i -can t events dur ing shu tdown are presented in Table - 4.


61/233

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ER 13227-8


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CONFIDENTIER 13227-8


64/233

III-10

T AB LE - 3Stage I Engine Performance Corrected to Standard InletConditions at 87FS. + 55 Seconds

ParameterThrust, engine (Ib)Specific impuse,engine (sec)Mixture ratio, engineOxidizer flow rate,overboard (Ib/ sec)Fuel flow rate,overboard (Ib/ sec)

AcceptanceTest*432,603261.421.95191093.87560.94

Predicted Flight(including 4850-lbthrust growth)*437,453261.421.95191106. 13567.22

FlightPerformance*434,328261.451.94211096.26564.97

*Martin~Baltimore modified thrust coefficient relationship used.TABLE III-4

Stage I Engine Shutdown ParametersParameter

Time from P decay to 87FS0 (sec) 2PC at 87FS2 (psia)Time from FS 9 to data dropout (sec)P at data dropout (psia)

S/ A 10.134400. 7145

S/A 20. 175250. 7105

e. Engine malfunction detection system (MDS)The Stage I engine MD S operat ed satisfactorily and within specifiedlimits throughout the flight. Figures - 1 and III -2 illustrate responsetimes and actuation pressure levels of the malfunction detection thrustchamber pressure switches (MDTCPS) during engine start for S/ A 1and S/A 2, respectively. F igures III-4 and III- 5 show deactuationtimes and pressure levels during shutdown for S/ A 1 and S/A 2, r e-spectively.A summary of the operating characteristics of the switches is tabu-lated in T ab le - 5 .

1 1 1 I l l i l ^ ,ER 13227-8


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Ill-11TABLE - 5

Stage I MDTCPS Operation

1

ActuationTime (sec)FSX + 0. 945FS . +0.935

Pressure*(psia)600580

DeactuationTime (sec)FS - 0.035FS 2 - 0.025

Specification RequirementsActuation 540 to 600 psiaDeactuation 585 to 515 psia

Pressure' tpsia)585535

g rate of the MDTCPS actuation times precludes accuratepressure.f. Engine prelaun ch malfunction detection system (PMDS)All PMDS switches actuated within the specified times and pressuresshown in Table ~ 6 . Ho we ve r, the oxidizer pressurant pressure

(OPPS) cycled once, prior to final actuation at MOCS TQ + 1. 67487FS, + 1.623 seconds) at a pressure of 414 psia. Approxi-7. 5 milliseconds after initial actuation, the switch was deactivatedmilliseconds before final "make, " as shown in Fig. III-6. Th e"make" pressure (414 psia) agreed closely with the prefiight"make" pressure (410 psia).

TABLE III-6Stage I PMDS Operation

timeMeasured time from 87FS. (sec)Measured time from Tn (sec)Required time (sec)*

pressureMeasur ed (psia)Required (psia)

TCPS

1.0121.063T0+ 2. 2

**600 to 640

OPPS

1.6231.674TQ+ 2. 2

414360 to 445

FPDSP

0.9020.953T Q +2 . 2

j j c s f e46 to 79psidshutdown timers start from TQ; 87FSt is 70 to 100 milli-seconds after TQ.

Not instrumented.

ER 13227-8 HH.


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ER 13227-8 ! i

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67/233

CONFIDENTIAL - uA review of oxidizer pressurant orifice inlet pressure (POPOI)failed to disclose f l uc tua t i o ns which may have contributed to the cycling

of the switch. The slow sampling rate of POPOI (40 samples/second)makes observation of such pressure fluctuations di f f icu l t . There is nohistory o f pressure f luctuat ions during POPOI bu i ldup o n previous G L Vflights or during engine acceptance tests. It is highly improbable,therefore, that switch cycling resulted f rom pressure changes.A n o t h e r possible cause of switch oscillations may be sensitivity ofthe switch to the vibrational environment w h e n pressure is held steadyat or near the actuation level of the switch. A test history for thissituation does not exist; therefore, a firm evaluation of this cause isn o t possible at the present time. Implementation of a test program to

verify the validity of this assumption is u n d e r consideration.An indication o f switch oscillation could be the result of the genera-t ion of a spurious electrical signal in the landline circuit. However,this assumption does not appear valid since the instrumentation systemshows n o evidence o f a transient.Prior to the start of the interrogation period, at M O C S Tn + 2. 2

seconds, chattering of the OPPS has no ef fec t upon engine operationan d c a n n o t p r o d u c e an engine s h u t d o w n . OPPS chatter after initiationof interrogation can result in an engine shu tdown if the duration o fs w i t c h d ro p o u t exceeds 40 milliseconds. This time interval is neces-sary for "make" and latching of the holdfire monitoringrelay ( H F M R ) ,which wou ld then initiate a s h u t d o w n .T h e very short duration o f OPPS d r o po u t (7. 5 milliseconds) o nG L V - 8 and its occurrence prior to M O C S Tn + 2. 2 seconds precludedan inadvertent engine s h u t d o w n .

2. Stage I I E n g i n e ( Y L R 9 1 - A J - 7 S / N 2009)a. C o n f i g u r a t i o n and special proceduresT h e G L V - 8 Stage engine was d i f f e r en t f rom that o f G L V - 7 in twoareas. The lockwire on the turbine rotor retaining bolts was replacedwi th a lockring for increased bolt security (ECP- 059). Secondly, theconf igura t ion of the thrust chamber injector was changed (ECP- 157).T h e new G E M S I P injector (effect ive for G L V - 8 and up) is less susceptiblet o combus t ion instability and provides essentially the same performanceas the replaced pr o d u c t i o n q ua d l e t injector.T h e propulsion system break- of- integrity review showed that the

r e d u n d a n t engine shu tdown valve had been removed at Martin- Baltimorean d that retest procedures were not adequate to verify that closures


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I I I - 1 4

w e r e n o t inadver tent ly le f t in d u r i n g reassembly. T h e valve was re-moved a t ETR fo r inspect ion an d closures we r e n o t found in the un i t .As on Stage I , the th rus t cham ber valve th rough - bo lt s were re to rqu edto th e f inal 300 to 320 i n c h - p o u n d s at E T R .b. Start t r ans ien tT h e engine start t r a ns i e n t , as i l lustrated by the t h r u s t c h a m b erpressure his to ry in Fig. I I I -7 , showed an abnormally slow P rise rate

f rom th rus t ch am ber ign i t ion t ime to stagin g blackou t . F ol lowing te lem -etry b l a c ko u t , no r ma l P opera t ion w as observed. T h e slow P riserate i s no t indicat ive of engine operat ion , since th e p u m p dischargepressures, turbine speed, an d M D F JP S a c tua t io n e xh i b it e d n o r m a lcharacteristics. A similar phenomenon was observed o n G L V - 5 . T h emo s t likely cau se o f the G L V - 5 an d G L V - 8 occurrences was moisturef r e e z i n g in the P sen sing l ine or t ran sdu ce r cavity d u rin g Stage I fl ight .Othe r possibilities inc lude con taminat ion or an improper ly sized trans-ducer or i f ice .

Significant engine start e ve n t s are presented in Table I I I -7 .TABLE - 7

Stage Engine Star t ParametersParameter

F S* to ini t ial P rise (sec)C3P ignition spike (psia)C3P step (psia)C3P overshoot (psia)C3

S / A 30 . 6 5InvalidInvalidN ot available*

*Staging blackout period,c . Steady- st a te pe r fo rm an ceStage en gin e s teady- s ta te fl ight p e r f o r m a n c e was sat is factoryt h r o u g h o u t f l ight . T h e average Stage I I engine pe r fo rmance in tegra tedover steady- state operat ion (from F S . + 1.2 seconds to 91FS2) agreedclosely with pref l ight predic t ion s. Average f ligh t perfor m an ce isc o mp a r e d in Table III-8.


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1000CONFIDENTIAL

- '

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+3.0

Fig. III-7. s/A 3 Start Transient

CONFIDENTIER 13227-8


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- 16

TABLE III- 8Predicted and Flight Performance Comparison- -Stage II Engine

ParameterThrust, chamber (Ib)Specific impulse, engine (sec)Mixture ratio, engineOxidizer flow rate,overboard (Ib/sec)Fuel f low rate, overboard(Ib/sec)

PreflightPredictedAverage101,750311.841.7680208.57117.72

FlightAverage101,820311.381.7901209.96117.04

Difference (%)+0.07-0. 15+ 1.25+0.67-0.58

The engine flight performance calculated with the Martin- BaltimorePRESTO computer program is shown in Fig. - 8 as a function of timefrom 91FS.. The preflight prediction is presented for comparison.Flight performance (Fig. -8) shows unusual operation of theStage II engine beginning at appr oximately 91FS, + 90 seconds. A

decay in performance and an unusual time- dependent characteristicare predominant in oxidizer flow rate, mixture ratio, and thrust,amination of the data indicat es that th e oxidizer pump performance wasabnormal after 91FS* +90 seconds. After this time, oxidizer and fuelpump discharge pressures, thrust chamber pressure, and turbinespeed are normally expected to increase. At approximately 91FS9+ 90 seconds, P and P started to decline. The pressure declineod 3 c3lasted approximately 30 seconds and then erratically followed the ex-pected trend (increase) but at the lower level. Turbine speed and P f.ta3remained essentially constant during this period; however, P- H showedsome respon se to the fluctuations in P and Pod 3 c3

Thrust chamber pressure also indicated larger pressure oscillationsthan noted on previous GLV flights, beginn ing at approximately 91FS1+ 140 seconds and continuing until the end of Stage flight. Prior tothis time, the P pressure fluctuations were approximately+15 psi,C3increasing to a maximum of approximately +40 psi near 91FS. + 150seconds. These steady-state abnormalities are under investigation.

E R 13227-8


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III -17

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JS load (Stage I I

R e s u m eloadLo a d jc o m p i

M e a s 4432(Stage I I f l o w m e t e r ) |M e a s 4431(Stage I f l o w m e t e r )M e a s 4425 ( R S V ) -

0300 0400Eastern Standard T ime (hr)

Load ing i jcomplete I j

Fig. HI-12. Fuel Temperature D uring Loading

ER 13227- 8


87/233


88/233

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30 40 45 50Bulk Fuel Temperature (" F)

Fig. H X - l A . Propellent Bulk Temperatures at Liftoff, Stage II

CONFIDENTIALE R 13227 -8


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Fig. 111-16. Stage I Fuel Tank Bottom Probe Temperature (Meas

QjliiriPn III'IER 13227- 8


91/233

- 37

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Fig. 111-17- Stage Oxidlzer Tank Bottom Probe Temperature (Meas 4604)

NFIDENTIALER 13227- 8


92/233

- 38

45

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0400 0800Eastern Standard Time (hr)

1200

Fig. m-18. Stage II Fuel Tank Bottom Probe Temperature (Meas 4601)

CONFIDENTIALE R 13227- 8


93/233

.CONFIDENTIAL m - 3 9opened. I t is indicated that higher- than - predicted heat t ran sfer coef-ficients are incurred in the Stage II fuel tank when the curtains areopened. Changes to the propellant temp erature p rediction p rogramwill be made to cover this situation, thereby making erector curtaindeployment a more predictable variable when forecasting propellanttemperatures.

The lack of correlation of the tank bottom probe readings with pre-dictions after the curtains were opened was also similar to the G L V- 1experience. Analysis of previous flights shows that the thermalstratification experienced on G L V - 1an d G LV- 8 was as much as 15%greater than predicted. Table - 20 shows the predicted an d actualstratification for G LV- 1, G LV- 5 and G LV- 8; G LV- 5 is typical of the"no curtains opened" operation.TABLE III- 20

Thermal Stratification- - Predicted and ActualSystem

Stage IOxidizerFuelStage IIOxidizerFuel

Predicted

25%5%10%15%

G L V - 1

31%18%19%32%

G LV- 5

23%5%10%19%

G LV- 8

39%15%17%30%

It is believed that a large increase in stratification occurred whenthe Stages I and oxidizer curtains were opened at 0700 EST. Thiswould explain why the probe tem perature rise rate failed to respon dto the wind increase which occurred coincident with the curtain open-ing. Changes in the rise rate program to cover this situation will alsobe made for G LV- 9 and up.d. Suction temperaturesThe actual pump inlet temperatures were in good agreement withthe predicted temperature profiles. These data are shown in Figs.Ill- 19 through - 22. The trends of the actual temperature curvesare in fair agreement with those predicted. D eviations may beascribed to differences in predicted and actual weather and the dif-ferences between predicted and actual thermal stratification . Thermalstratification differences (predicted versus actual), while large forpropellant temperature monitoring uses, are not reflected as large ona propellant temperature profile basis.

ONFIDENTIAL .ER 13227- 8


94/233

- 40

! F-45 day predicted- -: M eas 0024 j Tank bot tom probe ABest es t ima te

80 100 120 140 160Time from 87 FS1 (sec)

Fig. 111-19. Stage I Oxidizer Pump Inlet Temperature (Meas 0023 and 002 )

13227-8


95/233

CONFIDENTIAL - 41

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- 52

TABLE III- 25Best Estimate Average Sensor Uncover Time

QuadrantII/IV / II I II III I IIVI / I VI I / I I IIII IIVI IIV

Meas00560058005900540052005300500060054205480549054005460547054505500544

0551

SystemStage I oxidizerStage I fuelStage II oxidizerStage II fuel

SensorStage I oxidizer highStage I oxidizer outagesStage I oxidizer ou tage/Stage I fuel highStage I fuel outage - jStage I fuel outage /Stage I fuel shutdown )Stage I fuel shutdown /Stage II oxidizer highStage II oxidizer outage )Stage II oxidizer ou tage/Stage II fuel highStage II f u e l outage )Stage II fuel outage /Stage II oxidizer shutdown)Stage II oxidizer shutdown/Stage II fuel shutdown

Stage II fuel shutdown

Integrated AverageTemperature(between uncoverings) ( F)42.843.444. 142. 1

AverageUncoverTime ( G M T )1641:13. 214)1643:34.338)1641:15.314)1643:30. 388 )

1643:36. 1881644:35. 563)1646:31. 489)1643:54.013)1646:35.014)

1646:38.989Did notuncoverDid notuncover

At(sec)

141. 124

135.074

115.920

161.001

Corresponding D ensity(lb/ ft 3)92. 17057.34192.06857.220

IDENT IALER 13227- 8


107/233

NFIDENTIAL - 53

4. Propellant Utilizationa. Level sensor uncoverFigures - 27 and 111-28 show the predicted, actual and reconstructedlevel sensor uncover times for Stages I and II. Measured level sensoruncover times are tabulated in Table 111- 24.Slosh, as indicated by on and off signals at the time of level sensoruncovering, was minimal on this flight. All sensor uncoverings wereclean.b. Best estimate level sensor uncover timesTable 111-25 contains the best estimate average level sensor uncovertimes for the G T - 8 flight. Also shown are the integrated averagetemperatures between level sensor uncoverings and the correspondingdensities. The measured average uncover times shown in Table 111-24were dec reased by 0. 058 second to allow for the built- in level sensordelay of 0. 033 second and for the PCM digital sampling rate of 0. 05second.Table >26 contains the level sensor volumes and delta volum esused in the level sensor flow rate analysis. The Stages I and IIoxidizer and fuel high level sensor volumes were reconstructed toreflect the volumes which were determined by calibration at CapeKennedy using the propellant t ran sfer and pressurization system.The Stages I and II outages and shutdown level sensor volumes werecalculated using the actual counts of flowmeter pulses during thespecial loading.

TABLE III- 26Averaged Volumes at Level Sensor Locations

TankStage IoxidizerStage IfuelStage IIoxidizerStage IIfuel

SensorH igh- levelOutageH igh- levelOutageH igh- levelOutageH igh- levelOutage

Averaged Volumes(stretch included)(ft 3)

1710.0737.80

1403.0066.02

286.342 2 . ( i ! j

349.3217.76

DeltaVolumes(ft 3)

1672.27

1336.98

263.65

331. 56

NFIDENTIAUER 13227- 8


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- 56

. Flow ratesTable - 27 presents the predicted and the actual volumetric f l ow

rates between level sensors.TABLE III - 27

Propellant Volumetric Flow Rate

T a n kStage I oxidizerStage I f u e lStage II oxidizerStage I I f u e l

Predicted(ft3 / sec)10. 52910. 067

2. 0972. 115

Actual( ft 3/ sec)11. 8509. 898

2 . 2742. 059

d . Mixture ratioTable - 28 shows the Stages I and II predicted and actual in- flightaverage engine mixture ratios for G LV- 8.

TABLE III - 28Engine Mixture Ratio

SystemStage IStage I I

Mixture RatioPredicted1. 94431. 7680

Actual1 . 9 2 9 01. 7901

Sensitivity coefficien ts applied to the delta between the predicteda n d actual variations in average suction pressure and temperaturebetween sensor uncoverings yield the information shown in Table 111- 29.TABLE I I I - 2 9

Mixture Ratio Pressure and Temperature

SystemStage IOxidizerFuel

APressure(psi)

. + 1. 8Total Stage I

AMixtureR a t i o(press. )

- 0. 0- 0.006336- 0. 006336

ATemper-ature ( F)

- 1. 3- 0 . 7

A M i xt u r eR a t i o( t e m p )+ 0. 003013- 0.001139+ 0. 001874

AMixtureRatio(total)

+ 0. 003013- 0. 007475- 0. 0 0 4 4 6 2

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IFIDENTIAL - 57TABLE 1- 29 (continued)

SystemStage IIOxidize Fuel

APressure(psi)

0.0- 2. 0Total Stage II

AMixtureRatio(press. )

+0.0+0. 008850+0. 008850

-ature ( F)- 2.0+1.9

MixtureRatio(temp)

+0. 005230+0.003162+0.008392

AMixtureRatio(total)

+0.005230+0. 012012+0.017242

By applying the delta mixture ratio (total) shown in Table - 29 tothe predicted (F - 45 day) between- sensors mixture ratios, the run- to-run variation can be calculated. The mixture ratio deviation alongwith the allowable run- to- run dispersions is shown in Table - 30.TABLE III-30

Mixture R atio D eviation

SystemStage IStage II

PredictedM ixture Ratio*1.93981.7852

ActualMixtureRatio1.92901.7901

D eviation

- 0.56+0.27

AllowableRun - to -RunD ispersion

+1.38+2.28

Corrected for pressure and temperature variations.e. Outage and trapped propellantTable 111-31 shows the mean and maximum (99%) outages predictedfor G LV- 8. Also shown are the actual outages as calculated using theinformation contained in the reconstructed propellant inventories ofTables 111-37 and - 38.

TABLE - 31Outage Prediction

SystemStage I

Predicted (F - 45 day)Mean

0.220%566 Ib

Max (99%)0.643%1652 Ib

Predicted (F - 0 day)Mean

0.315%810 Ib

Max (99%)0.744%1910 Ib

Actual0.098%oxidizer252 Ib

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- 58 CONFIDENTIAL

TABLE - 31 (continued)

SystemStage I I

Predicted (F- 45 day)M e a n

0. 344%206 IbM ax (99%)

1 .026%615 Ib

Predicted (F- 0 day)M e a n

0. 517%310 IbM ax (99%)

1.217%730 IbActual0. 436%fuel261 Ib

All outages are presented as percent of total steady- state propellants(taken from Ref. 11) and in pounds. The values used for total steady-state propellants are: 256,737 pounds for Stage I and 59,966 poundsfo r Stage II.T h e trapped propellants for Stages I and II are given in Table - 32.

T A B L E 111-32Trapped Propellants

SystemStage IAbove interfaceBelow interfaceStage IIAbove interface

Below interface

Oxidizer (Ib)0235020

Fuel (Ib)20309

014f. Start and propellant consumptionsT h e predicted and actual propellant consumptions during the Stage Istart period are sh own in Table III- 33.

TABLE - 33Stage I Ignition and Propellant Consumption

Stage IStart consumption( 8 7 F S J to TCPS)Holddown consumption( T C P S to liftoff)

Oxidizer (Ib)Predicted

2092157

Actual209

2218

Fuel (Ib)Predicted

441128

Actual44

1184

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CONFIDENTIAL - 59T h e predicted and actual start consumptions listed in Table 111-33were selected from Ref. 15 and were modi f i ed to allow for the d i f f e r e n cebe t w e e n propellant out of the tanks (as listed in the report) and pro-pellant overboard. The predicted holddown consumption was derivedf rom the engine analytical model and previous flight test data, whilethe actual value was derived from the Post- test Rocke t Engine System

Total Operation (PRESTO) engine performance reconstruction program.The Stage propellant consumptions between 91FS. and 91FS1

+ 1.2 seconds are listed in Table -34.T A B L E - 34

Stage II Start Propellant Consumption

Start consumption(91FS to 91FS + 1 . 2 sec)

Oxidizer (Ib)

135

Fuel (Ib)

53T h e consumptions were obtained from Ref. 15 and modified as on theStage I start consumption.

g. Vapor retainedT h e predicted and actual values of vapor retained in the tanks as aresult of pressurization gases and propellant vaporization during flightare shown in Table - 35.

T A B L E - 35Pressurization Gas Inventory

SystemStage IVapor retained:

Oxidize r tankFuel tankVaporizedStage IIPressurization

Fuel tankVaporization

Oxidizer tank

Oxidizer (Ib)Predicted

31886

5

9

Actual*

32286

4

9

Fuel (Ib)Predicted

0900

50

-

Actual*

0910

49

-*Actual values were obtained from reconstructed flight pressure profileof pressurization computer program runs.

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CONFIDENTIAL - 63

TABLE III- 39Prevalve Identification

D escriptionStage I oxidizer S/ A 1(fill and drain)Stage I oxidizer S/ A 2(drain)Stage II oxidizer S/ A 3(fill and drain)Stage I fuel S/A 1(fill and drain)Stage I fuel S/A 2(drain)Stage II fuel S/A 3(fill and drain)

Part N o.PS47510007- 139PS47510007- 159PS47510005- 199PS47510005- 159

PS47510005- 169PS47510006- 059

Serial N o.0700025070002906000220600023

06000990400013

b. Level sensorsG LV- 8 incorporated 18 Bendix optical type propellant level sensors.These are identified in Table III - 40. All sensors performed satisfactorilyduring propellant loadings and in flight.c. Oxidizer standpipeThe oxidizer suct ion line standpipes were charged with the remotecharge system at - 59 minutes. N o problems were encountered duringthe charging operation. F light data obtained from pressure Meas 0033and 0034, located in the standpipes, show surge chamber performanceto be normal and consistent with the low longitudinal oscillatory levelsexperienced on this flight.d. Fuel accumulatorsAccumulator piston response on this flight was similar to that ofother flights utilizing the same configuration (G LV- 3 through G LV- 6).This response is presented in Fig. HI-29.Dynamic friction levels for dry accumulators were measuredprior to installation of accumulator assemblies at M artin- Baltimoreand again prior to flight at ETR. A summary of these friction meas-urements is presented in Table 111-41 as peak- to- peak values (twicethe equivalent friction force in one direction).

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f

a>'na" *H'"uadrant1a'o38*I1ocation

to

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1TI

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I

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I

tC

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C

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-31 ^fcuD4)

6 Q.(yaw-right _error)

Primary system yaw error (Meas 0767)EHIG S yaw error (Meas 0744)

- 2 . 0 - (yaw-left!_rate)

Primary system roll error (Meas 0768)... ;:! U i

IG S roll error (Meas 0 7 4 5 )

(roll-CCWerror)155 165 175 185Time from Liftoff (sec)

Fig. VI-3. Stage II IGS Ya w / R o l l Flight History

ER 13227-b


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V I - 7

The data s h o w n in Table VI-2 indicate that the S E C O time delay f r o mg r o u n d station issuance to 91FS9 was 52 8 milliseconds. The timedelay b e t ween 91FS2 a n d A S C O reception wa s 4 4 3 0 milliseconds.

B . S P A C E C R A F T I N E R T I A L G U I D A N C ES Y S TE M A S C E N T P E R F OR M A N C E1. Prelaunch Nulls

The prelaunch IGSattitude error null signals were as follows:Pitch -0. 031 degreeYaw -0 . 149 degreeRoll +0. 079 degree

These signals were well within the specification values of 0 . 37 degreein pitch and yaw and 0 . 2 5 degree in roll.2 . Stage I Performance

IG S performance during Stage I flight correlated well with the primarysystem, as s how n by a comparison of IGS and corresponding primarysystem attitude errors in Figs. IV-2 through IV-4. The dispersionsb e t ween IGS and primary system attitude errors at BECOare discussedin Chapter IV .The IGS Stage I gain change discrete was issued at LO + 105.031seconds 0 . 025 second, which w a s well within th e specification time o f104. 96 seconds, 1%.

3. Stage II PerformanceIG S pitch, yaw and roll performance during Stage II Flight was normal.The attitude error dispersions which had built up between the IGS andprimary system during Stage I flight in pitch, yaw and roll were apparentin the early portion of Stage II flight, as s how n in Figs. VI-2 and VI-3.a. Stage II pitchIG S Stage II pitch attitude error appears in Fig. VI-2. Primary systempitch attitude error and RGS pitch steering commands are s h

Launch Vehicle No. 8 Flight Evaluation

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