Launch Vehicle No. 7 Flight Evaluation

8/8/2019 Launch Vehicle No. 7 Flight Evaluation

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( N A S A - C F - 8 3 0 8 8 ) L A U N C HV E H I C L EN O . 7 F L I G H TE V A L U A T I O N( M a r t i n Co.) 2 8 9 p

i -MSC-G-R-66-1Supplemental Report2January 1966

d as: S upp lemental Report 2To: GeminiProgram Mission Report

GeminiVIEMSC-G-R-66-1

By: G emini V n Mission E valuation T eamNational Aeronautics and Space

AdministrationManned Spacecraft CenterH ouston, Texas

F O R N A S .

P E R S O N N E LO N L Y

LAUNCHrVEHICLE NO. 7t

FLIGHT

EVALUAT ION (U)

U . S. G o v e r n m e n t A g e n c ie s

January 1966

I T E DS T AT E SA IRF O R C E. . . L osAngeles, California&


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E R 1 3 227 -7 J a n u a r y1966

G E M I N IL A U N C HV E H I C L E

N A S A - M S C - G - R - 6 6 - 1Supplemental Report 2January 1966

LAUNCH VEH I CLE NO. 7

FLIGHT

EVALUATION (U)

Approved by

A F F E C T L .S TAT E SWITH.. . _L AW S ,HUE 18, U.TRANSMISSIONORIN AMY MANNE"PROHIBITED

D O C U M E N T.W '

_,.J INFORMATION-EN'SEO F T H EU N I T E D O F T H EESPIONAGE

,743 AN C/ 9 4 . IT S

i tCL. J. Rose

Ass i s t an t Technical Director

Test Evaluation

Issued as: Supplemental Report 2To: G emini Program Mission Report

Gemin i VI IMSC- G- R- 66- 1

Prepared by

MARTIN COM PANY, BALTIM ORE DIVISIONBaltimore, Maryland 21203

Under C O N T R A C TAF 04(695)- 394P R I O R I T YDX- A2

I. C. Curlander

Technical Director

By: G emini Mission Evaluation TeamNational Aeronautics and Space

Ad m inistrationM a n n e d Spacecraft CenterHouston, Texas

Down

Fo r inteafti

S PA C E S Y S T E M SDIVISIOA IR F O R C ES YS T E M SCOMMAND

U N I T E DS TAT E SA IR FORCELo s Angeles, California

year


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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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CONTENTS

Page

Foreword u

Summary v ii

I . Introduction I -1

II. System Performance II-1

A . Trajectory Analysis II-1

B. Pay load Capability 11-39

C. Staging 11-39

D. Weight Statement 11-39

III. Propulsion System III-l

A . Engine Subsystem III-l

B. Propellant Subsystem 111-22

C . Pressurization Subsystem 111-67

D. Environmental Control 111-78

IV . Flight Control System IV-1

A . Stage I Flight IV-1

B . Stage I I Flight IV-8

C. Post-SECO Flight IV-13

V . Hydraulic System V - l

A . Stage I V- l

B . Stage I I V-5

V I . Guidance Systems VI-1

A . Radio G u i d a n c e System Performance VI-1

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L\

CONTENTS (continued)

Page

B. Spacecraft Inertial Guidance System AscentPerformance VI-5

VII. Electrical System Analysis VII-1

A. Configuration VII-1

B. Countdown and Flight Performance VII-1

VIII. Instrumentation System VIII-1

A, Airborne Instrumentation VIII-1B. Landline Instrumentation VIII-2

IX. Range Safety and Ordnance IX-1

A. Command Control Receivers IX-1

B. MISTRAM IX-2

C. Ordnance IX-2

X. Malfunction Detection System X-lA. Configuration X-l

B. System Performance X-2

XI. Crew Safety XI-1

A. Prelaunch Winds Flight Simulations XI-1

B. Slow Malfunctioning Monitoring XI-6

XII. Airfram e System XII-1

A. Structural Loads XII-1

B. POGO XII-19

C. Recovered Stage I Oxidizer Tank XII-27

XIII. AGE and Facilities XIII-1

A. Mechanical AGE . XIII-1

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vii

S U M M A RY

On 4 December 1965, Gemini-Titan No. 7 (GT-7) was launched suc-cessfully and on schedule from Complex 19, Cape Kennedy, Florida.

Launch vehicle/spacecraft separation was completed 368.7 secondsafter l i f toff . Spacecraft re-entry was accomplished afte r completionof 13.8 days in orbit.

The 240-minute countdown was picked up at 1030 EST on 4 Decemberand continued without incident through liftoff at 1430 hours EST. Thespacecraft was inserted into an elliptical orbit with a perigee of 87nautical miles and an apogee of 177. 1 nautical miles; all test objectivesfor the launch were achieved.

Stages I and II engines operated satisfactorily throughout poweredflight. Stage I burning time was 159. 121 seconds, w i t h s h u t d o w n ini-tiated by oxidizer exhaustion. Stage II engine operation was terminatedby a guidance command after 181.403 seconds of operation.

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 neverexceeded 1. 7 deg/sec, and the maximum attitude error was 1. 1 degrees.The maximum rate and attitude error that occurred during staging didnot exceed 2.9 deg/sec and 2. 1 degrees, respectively.

The radio guidance system ( R G S )performa nce was satisfactory.Pitch and yaw steering signals and SECO discrete 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 launch opera-tions and both stages of flight. There wer e no significant pr essu reperturbations at liftoff or during flight.

The electrical system functioned as designed throughout the launchcountdown and flight. Power transfer to vehicle batteries was smooth.

All channels of the PCM instrumentation system functioned satis-factorily throughout the flight, resulting in 100% data acquisition. Thelandline instrumentation system also functioned satisfactorily prior toand up to liftoff. All airborne instrumentation hold functions monitoredin the blockhouse remained within specification throughout the count-down.

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Vlll

T he ordnance system umbilical dropweight release, propulsion sys-tem prevalves, explosive launch nuts and stage separation nuts operatedas designed. The performances of the command control receivers andthe MISTRAMtransponder were satisfactory.

Malfunctiondetection system (MDS) performance during preflight

checkout and flight was satisfactory. There were no switchover com-mands during the flight.

The flight environment encountered by GT-7 was within design re-quirements. Flight loads were well within the launch vehicle 1 s struc-tural capabilities. The most critical loading (which occurred at pre-B E C O , aft of Station 320) reached 98. 5% ofdesign limit load.

T he longitudinal oscillation ( P O G O )on GT-7 reached a maximumvalue at Station 280 of 0.125 g zero-to-peak at a frequency of 11.8 cpsa t L O +133. 3 seconds. This was the lowest P O G Oexper ienced on anyGemini flight to date.

Crew safety monitoring, which w as conducted at N A S A - M S C ,w asactive during prelaunch and the launch. All guidance monitor parameterswere nominal and no corrective action was required during the flight.

T he precount operation progressed without problems for the launch.All AGE and facilities operated without incident during the countdown.Propellant loading was completed within the scheduled time span andto the specified load and temperature limits.

T wo electrical umbilicals, 2B1E and 2B2E, disconnected out ofsequence by 0. 015 second; however, this is not detrimental to anysystem. Engine blast and heat damage to the launch stand was minor.

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I X

GLV-7 Test Objectives and Results

Objective

Primary

P-l Demonstrate satisfactoryboost by the launch ve-hicle system of a mannedGemini spacecraft intothe prescribed orbital

insertion conditions.

P-2 Evaluate launch vehiclesubsystem performanceduring powered flightfor mission success andcrew safety.

Secondary

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

S-2 Demonstrate ability toload propellants to theweight and tempera-ture limits imposed bypayload and vehicle re-quirements.

Results

P-l Orbit insertion was withinthe predicted tolerance forV, h and V.

P~2 All systems performedsatisfactorily throughoutflight. The POGO oscilla-tion (0. 125 g zero-to-peak)was the lowest encountered.

S-l Vehicle flight was withinthe 3 ~sigma predicted tra-jectory.

S~2 Tanks were loaded withinthe required tolerances ofweight and temperature.

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1-1

I. INTRODUCTION

This report presents an engineering evaluation of Gemini LaunchVehicle No. 7 (GLV-7) systems performance during the countdown,launch and powered flight phase of the Gemini 7 mission.

The Gemini-Titan No. 7 (GT-7) vehicle was launched on schedulefrom Complex 19, Cape Kennedy, Florida at 1430 hours EST on 4December 1965.

Gemini 7 was the sixth mission and the fourth manned flight of theprogram, with astronauts Frank Borman and James Lovell aboard thespacecraft. The 14-day mission, which included a rendezvous withGemini 6, was completed successfully on 17 December 1965.

The GT-7 vehicle was comprised of the two stage GLV-7 (similarto GLV-5) and the Gemini 7 spacecraft. The spacecraft was injected

into an elliptical orbit having a perigee of 87 nautical miles and anapogee of 177.1 nautical miles.

Significant events and tests for GLV-7 at ETR are summarized inFig. 1-1.

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G LV- 7 on dock, ETR

Erection of G LV- 7

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^0Jo in t combined sys tems t e tL V propellant t a n k i n g te s Jpacecra f t mechanica l m a1

-imulated flight test (SFT)

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I I -1

. SYSTEM PERFORM ANCE

A. TRAJECTORY ANALYSIS

1. Orbit InsertionG emini Laun ch Veh icle No. 7 (G LV- 7) performed as predicted an d

inserted the Gemini 7 spacecraft into earth orbit well wit h in the allow-able tolerance l imits .

A comparison of the predicted an d observed insert ion condit ions isgiven in Table II- l. In this table and in all succeeding references to apredicted (nominal) trajectory, the data have been obtained from theG LV- 7 45- day prelaunch report (Ref. 16), updated to reflect a space-craft weight of 8069 pounds (liftoff spacecraft weight- - 8085 pou nd s), T- lho ur win d an d atmosphere data, an d the - 1. 64% pitch an d - 1. 69% roll

programmer biases. The observed tra jecto ry paramete rs are thosederived by the M artin Compan y fro m the final G E Mod III- G 10 ppsr ad a r data. These data have been smoothed an d corrected fo r both re -fraction errors an d systematic biases by the G eneral Electric Corpora-tion before submittal to the Martin Company.

TABLE - l

Comparison of Insertion Con dition s at SECO + 20 Second s

Altit ud e(naut mi)

Inertia! ve-locity (fps)

Inertial flightpath angle(deg)

PlannedNominal

87.106

25,806

- 0.0004

GE

Mod - G

87.183

25,789

0.0765

ObservedMinus

Planned

+ 0.077

- 17

- 0.0769

Prel iminaryTolerance

+ 0.346

+ 30.38

+ 0.1251

2. Derivation of Trajectory Uncertainties

The expected maximum vehicle dispersions and RG S dispersionsat BECO an d at SECO + 20 seconds were obtained from R efs. 11 and12, respectively. A root sum square (RSS) of these dispersion s ist e rmed the preliminary tolerance. Afte r determ ination of the prelimi-nary tolerance, the total tolerance m ay be computed by the ari thm etic

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II-2

addition of the preliminary tolerance to the 3-sigma data error of theinstrumentation source being considered. Thus,

I 2 9Preliminary tolerance = V(vehicle dispersions) + (RGS dispersions)

Total tolerance = preliminary tolerance + 3-sigma data error.

The resulting preliminary tolerance is shown in Table II-2. Becausethe actual insertion conditions were within the preliminary tolerance,the data error estimates are not needed and, therefore, have been ex-cluded from this report.

3. Geodetic and Weather Parameters

Significant geodetic and weather parameters are shown in Table II-3.The winds were relatively strong, and atmospheric pressure and tem-perature were nearly standard. Winds were essentially sidewinds atvery low altitudes, shifting to westerly (almost pure tail wind) directionabove 3 0 0 0 feet.

4. Flight Plan

The primary objective for GLV-7 was to place the Gemini 7 space-craft into an elliptical earth orbit with an 87-nautical mile perigee* and183-nautical mile apogee. * Having achieved orbital insertion at 25, 806fps, ** the spacecraft then separates from Stage II (with no net changein velocity) and coasts to the desired apogee. The following flight planwas employed to attain these desired conditions.

A vertical rise is planned for the first 23.04 seconds following lift-off, during which time a programmed roll rate of 1.25 deg/sec is initi-ated to roll the vehicle from a pad orientation of 84. 908 degrees to theflight azimuth of 83.608 degrees.

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 as shown inTable II-4.

Guidance commands from the radio guidance system ( R G S ) are initi-ated at liftoff + 168.35 seconds and continue until two seconds prior to

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

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II-4

TABLE - 3

Geographic an d Weather Conditions at Launch

Location

Site

Site coo rdinat es:

Latitude (deg)Longitude (deg)

Pad orientation (deg)

Complex 19

28.507 N80.554 W

84.908 t rue azimuth

Weather

Am bien t pressure (psi)

Am bien t t empera tu re (F)

Dew point (P)

Relative humidity (%)

Surface win d:

Speed (fps)

Direction (deg)

Winds aloft (max):

Altitude (ft)

Speed (fps)

Direction (deg)

Cloud cover

14.736

68

72

16.9

340

43,500

176

258 t rue azimuth

0. 5 cum ulus, base at 16, 000 ft

Reference Coordinate System

Type

Origin

Positive X- axis

Positive Y- axis

P ositive Z- axis

Reference ellipsoid

M artin reference coordinate system

Center of launch ring, Complex 19

D ownran ge along flight azimuthtangent to ellipsoid

To left of flight azimuth tangent toellipsoid an d _ l_ X- axis

Forms a right- handed orthogonalsystem

Fischer

Launch

Initial flight azimuth (deg)

Roll program (deg)

83.608 true azimuth

1.3 cw

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II- 5

SECO; h owever , velocity cutoff com put ations con tinu e to SECO. Be-tween SECO an d SECO + 20 seconds, the engine shu td own impulse con-t inues to add velocity to the vehicle (approximately 84 fps), and thespacecraf t is considered to be separated from th e sustainer at SECO

+ 20 seconds, nominally.

TABLE II- 4

Planned G LV Pitch Program

Program

Step 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 com parison of the planned and actual sequen ces of events is con-tained in Table II- 5, and a profile of the G T- 7 fl ight superimp osed onthe range planning m ap appears in Fig. II- 1.

5. Trajectory Results

Analysis of the range data and M od III radar data indicates thatG LV- 7 liftoff was n o r m a l and the vehicle flew close to the prescr ibedascent t ra jec tory throughout Stages I an d II. The only significantdeviations in th e t rajectory occurred in th e first stage, where at BECO

t h e vehicle was 741 feet high.Table H- 6 contains a simplif ied reconstruction of the BECO condi-

tion s. This table lists the pr im ary fac tors con t r ibut ing to the p i tchplane t ra jec tory dispers ions at BECO an d sum m ar izes the effect ofeach. Alth ou gh the reco n struc ted BECO does n ot m a t c h the flight da taquite as well as for previous Gemini flights, the differences are wellwit h in allowable to leran ce limits.

In the yaw plane the flight did not deviate significantly from thepred icted tra ject or y; hen ce, a recon struction was considered un n eces-sary.

A compar ison of the predic ted nominal (with win d) t ra jec tory withflight resul ts is shown in Table II- 7. Inspectio n of the various radardata indicates tha t the Mod III, M ISTRAM and - band r a d a r s wereconsistent at BECO. At insertion, ho wever , M ISTRAM produced themost accura te results. This is verified by th e Berm uda ' data whic hproduced very similar values of velocity, altitude an d flight path angle.

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- 6

TABLE II- 5

G T- 7 Flight Events Summary

Measurement

0800/ 0801

F C B- 10

2104

0356

0 3 5 72 1 0 1

0 1 6 94 4 2 1

4 4 2 2

4423

0734

0734

0732

0 7 3 2

0 7 3 2

0728

0 7 3 2

0 7 3 20735

0741

0356

0357

0032

05020169

0855

0732

0740

0 7 5 5 / 0 7 5 6

07390 7 7 7

0519

0522

0 5 2 10799

0855

AB- 03

E v e n t

P o w e r transferM O C S T- 0

87FS (T- 0)

Stage I S/ A- 1 MDTCPS make

Stage I S/ A- 2 MDTCPS make

TCPS S/ A- 1 a n d S / A- 2

Launch nuts

First motion

S h u t d o w n lockout (backup)

Lif toff

Start roll program

E n d roll program

Start pitch program No. 1

Stop pitch program No. 1


F C S gain change N o . 1



Staging enable (TARS discrete)

IPS staging arm timer

Stage I S/ A- 1 MDTCPS break

Stage I S/ A- 2 MDTCPS brea k

87FS / 91FS (BECO)

Start P C. rise

Stage separation

Stage II MDFJPS m ake


R G S enable

First guidance command

Stage II shut down en ableGuidance SECO

91FS 2Shutdown valve relay

S h u t d o w n squibASCOStage II MDFJPS break

Spacecraft separation

G M T(hr- min- sec)

1928:34.6

1930:00.09

:00. 190

:01. 137

:01.277

:01.403

:03. 51

:03. 584

:03.601

:03.702

:23.05

:24.09

:26. 64

1931:31.74

:31.74

:53. 30

1932:02.36

:02.36

:27.85

:28. 86

:39. 268

:39.276

:39. 311

:39.956

:40. 01

:40. 001

:45. 72

:45. 69

:52. 04

1935:20.03:40. 704

:40. 714

:40. 711

:40. 742

:40. 767

:40. 869

1936:12.4

Time fromAct ua l

- 89. 1

- 3.61

- 3.512

- 2. 565

- 2 .425

- 2.299

- 0. 19

- 0. 118

- 0. 101

0

19.35

20.39

2 2 . 9 4

88.04

88.04

109. 60118.66

118.66

144. 15

145. 16

155. 566

155. 574

155. 609

156.254

156.31

156.299

162.02

161.99

168.34

316.33337.002

337.012

337.009

337.040

337.065

337. 1 67

368.7

Lif toff (sec)Planned

- 89.

- 3.43

- 3. 37

- 2 .2 7

- 2 . 2 7

- 2. 20

- 0.20

-0 . 10:0. 10

0

19.44

20.48

23.04

88.32

88.32

110.00

119.04

119.04

144. 64

145.00

155.51

155.51

155.57

156.22

156.30

156.47

162. 56

162.56

168.35

317.44338.43

338.45

338.47

338.47

338.48

338.75

358.45

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IAL - 7

TABLE II- 6

Reconstruction of G T- 7 BECO Con ditio ns

Measured Parameters

Thru st (+ 0. 264% = + 8 68 I b )

Wi n d (T- l hr)

Outage (45 lb)# (compa redto nominal)

PropeUant loading (+ 102 Ib)

Inert weight (+ 146 Ib)

Trend Indications

Pitch programmer e r ro r(- 1.46%)**

Pitch engine misalignment(+0.07 deg)

Specific impulse (+ 0. 253% + 0.70 sec)

Pitch gyro drift (+5. 0 deg/hr)

App are n t As (A + B)

Measured '

At(sec)

- 0. 25

- -

+ 0.28

+ 0.06

- -

- -

- -

+ 0.26

- -

+ 0.35

+ 0.044

Altit ud e(f t )

+ 530

- 2300

+ 7 6 5

- -

- 340

- 470

- 1030

+ 3 70

+ 7 00

- 1775

+ 733

AVelo cit y( fps)

+ 59

+ 41

+ 2

- 11

+ 4

+ 10

+ 19

- 7

+ 117

+ 145

(des)

+ 0. 10

- 0. 25

- 0. 02

- 0. 02

- -

- 0. 10

- 0. 25

+ 0.05

+ 0. 16

- 0 .33

- 0 .19

*Mean outage = 570 Ib , nom inal outage = 452 Ib

**Nominal = - 1.64% (bias)

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m -

respectively. Figures - 4 and - 5illustrate the deactuation timesa n d levels for S/ A 1 and Si A 2, respectively.

A summary of the operating characteristics of the switc hes is pre-sented in Table III-5.

TABLE -5

Stage I M D T C P S Operation

Swtch

S/ A 1

SI A 2

A c t u a t i o n

Time(sec)

FS. + 0. 945

FS. + 1. 085

Pressure(psia)

590

550

D e a c t u a t i o nTime(sec)

FS 2 - 0. 045

FS 0 - 0. 035A

Specif icat ion requirements:

Actua t ion 540 to 600 psia

D e a c t u a t i o n 585 to 515 psi a

Pressure(psia)

575

540

f. Engine prelaunch m a l f u n c t i o ndetection system (PMDS)

Al PMDS swtches actuated wi th in the specified actuation times andpressures as shown in Table III- 6. However, the OPPS actuated latert h a n o n previous Gemini flight.

T h e OPPS is used to mon itor the Stage I oxidizer aut ogenous systemoperation prior to release of the launch vehicle and to f u r n i s h a no-gosignal to AGE if (1) the switc h does not ac tua te by T n + 2. 2 seconds or

(2 ) th e switch deactu ates i n t he period from T Q + 2 . 2 seconds to TCPS

+ 1. 8 seconds.

O n G LV- 7 , th e OPPS actuated a t 87FS. + 2 . 022 seconds (T Q + 2 . 122seconds) as shown in Fig. - 6. The OPPS actuation times for G LVflights are tabulat ed in Table III-7.

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: - - : ! .; :

/ "" .i . .

j

^

..

.. .

i

.

. '

1

. .

* *

-_

-

....'

.

. ,

1

.

- 3 . 5

L I l l - 11

87FS, Time from T-0 (sec)

nniVlWl

Pig. HI- 6 S/A2 Start Transient

ER 13227- 7


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- 12

T AB L E - 6

Stage I P M D S Swtch Operation

Parameter

Actua t ion time

Measured (sec)

R e q u i r e d (sec)*

Actua t ion pressure

Measured (psia)

R e q u i r e d (psia)

P S

87FSj + 1 . 21 3

M O C S TO +1.313

+ 2 . 2

**

600 to 640

OPPS

87FSj + 2 . 0 22

MOCS TQ +2. 122

+ 2. 2

405

360 to 445

FPDPS

STF Sj + 0. 988

MOC S TQ +1.988

+ 2 .2

**

46 to 79 (psia)

*The shu tdown timers start from MOCS T Q ;87FS 1 is 70 to 100milliseconds after T n .

**Not instrumented.

TA B L E I I I - 7

Summary of all GLVOPPS A c t u a t i o n Times

Vehi cle

1

2

3

4

5

7

OPPS A c t u a t i o nTimeFrom 87FS. (sec)

1. 8171.675

1 . 6 2 51. 722

1. 7682. 022

From T-0 (sec)

1. 9421. 7751. 7171.816

1. 8682. 122

Average = 1 . 77 2 secon ds Avera ge = 1. 873 secondsf rom 87FSJ from T-0

Average + = 2 . 1 92seconds Avera ge + = 2 . 3 05seconds

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Ill-13

A detailed discussion of the engine start transient will be necessaryto fully explain the cause for the late actuation of the OPPS; however,in summary, the start cartridge performance has the greatest influ-ence on engine start transients. For a given engine, on a run-to-runbasis, the start cartridge burning rate and duration (which are differ-ent for every cartridge) determine the time and rate at which theengine bootstraps to steady-state conditions. T he G LV- 7 S / A 2 starttransient (for example, chamber pressure buildup in the bootstrapcorridor) w as well within the Titan I I / G LVexperience, b ut on thelower side of the G L Vexperience.

Factors, other than start cartridge performance, that affect theOPPS actuation time are:

(1) Pressure setting of the OPPS switch

(2) Pressure level within the autogenous system as governed bythe back pressure orifice diameter

(3) Rupture characteristics of the engine burst diaphragm

(4) Overall steady-state pressure level within the engine.

In addition to the late actuation of the OPPS, the oxidizer pressurantorifice inlet pressure (POPOI) remained within the specification limitsfor OPPS actuation (360 to 445 psia) until approximately Tn + 4 seconds,well beyond th e time of OPPS interrogation. H ad the switch actuatingpressure been on the high side of the band instead of at 405 psia, anautomatic shutdown signal would have been initiated.

Following the G LV- 7 flight, a change in the oxidizer pressurantback pressure orifice diameter from 0. 50inch to 0. 46inch was madeon G LV- 6 . This change provided increased confidence that the OPPSwould make within the critical time period and also increased thesteady-state level of POPOI by approximately 80 psia.

Investigation will continue in the area of start transients to betterdefine the corrective actions for G LV- 8 succeeding vehicles.

2. Stage I I Engine (YLR 91-AJ-7, S /N 2008)

a. Configuration and special procedures

T he GLV- 7 Stage I I engine configuration was identical to that ofG LV- 5 .

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- 14

b. Start transient

Evaluation of Stage II engine data (Fig. Ill- 7) ind icated a normalstart transient except for a disturbance in chamber pressure (P ) at

91FSj + 1. 26 seconds. The remainder of the start tra n sien t was

normal. The initiation of the disturbance was followed by the charac-teristic ringing of the P m easurem en t when inco rpo ratin g a CEC

C 3tra n sduc er. These oscillations are similar to those normally experi-enced following th e rap id c ham ber pressure rise of th e ignit ion spike.

The disturbance in P was not observed in any other engine para-C 3

meter. If the disturbance ha d been an actual indication of the pressureconditions in the com bustion chamber, the pressure disturbance wouldhave been tran smitted hydrau lically thro ugh the engine an d would havebeen observed in the pump discharge pressures. Th ere wer e no per-turbations in any flight control or hydraulic actuator parameters at thetime of the indicated P disturbance.

It is con cluded that the pressure disturba nc e was no t in the com-bustion chamber but was caused by ignition of propellants or vapors inthe P instrum ent ation line. Similar start transient pressure disturb-

ances occurred on Titan II flights N- 22, N- 24, N- 28 an d N- 29 with no

detrimental effect on the engine. Pressure disturbances have alsobeen observed by Aer o jet on grou nd tes ts during both the start transientand steady state.

Significant engine start event s are pr esen ted in Table III- 8.

TABLE III- 8

Stage II Engine Start Parameters

Paramete r

PS,

to initial P rise (sec)C 3

ignition spike (psia)

P step (psia)C 3

FlightPerformance

0.64

635

490

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1000

-

a-

-.-

i:.

-

^-;

U

MDFJPS (Meas 0855)

+ 2 . 0Time from 91FS. (sec)

__^ ^ | ^

+3.0

Fig. HI-7- S/A 3 Start Transient

ER 13227-7


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- 16 NTlAb

TABLE III- 8 (con tin ued )

P a r a m e t e rFlight

Per fo rmance

P overshoot (psia)

FS. to P disturbance (sec)C 3

P disturbance (psid)C 3

*Not available

1 26

187

*Staging blackout period.

c. Steady-state performance

Stage II engine s teady-s ta te fl ight performance was satisfactorythroughout fl ight an d agreed closely with preflight predictions. Theaverage Stage eng ine pe r fo rmance in tegra ted over st eady- s ta teoperation (from FS . + 1 .2 seconds to 91FS 9) is compared to the pre-

flight prediction in Table III-9.

TABLE - 9

Predicted an d Average Stage II Engine Performance

P a r a m e t e r

Thrust, cham ber (Ib)

Specific impulse, engine (sec)

Mixture ratio, engine

OxLdizer flow rate, overboard(Ib/sec)

Fuel flow rate, overboard(Ib/sec)

PreflightPred ic tedAverage

101,979

313.63

1. 7621

207.61

117. 55

FlightAverage

102, 888

313. 24

1. 7785

210. 41

118. 05

Difference(%)

+ 0. 89

- 0. 12

+ 0. 93

+1.35

+0.43

The engine flight performance calculated with th e Martin PRESTOprogram is shown in Fig. - 8 as a function of t ime from 91FS r The

preflight prediction is also presented for comparison.

_ illlllllUPIER 13227- 7


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^^^^ w*-17- I^ C O N F I D E N T I A L

IIb17

-:

--

D

-:

1-

-

.-I

:'

E125 ,-

(:

120

115

110

105

104

102

.

315

1- 5 g 310

?" 305

1.85

"- 180

!

100 I

- L )_ N

Xh

:

215,- 1W

210

205

200

195

190

1.75

1.70

-

1

- -

300

I

rt

98

96

94

S2

o

F c

sp

,

5 0~! G (

< S * >G G O G

O Q < D O O

91FS,

20 100 120 140 160 180 200Time from 91FS 1 (sec)

Average Engine Performan ce Integratedfrom First Steady-State to 91FS 9

Symbol

F c (lb

4 (sec)MR e

W oo ( lb/sec)

W fo ( lb /sec)

Pref l ightPredic t ion

101979

313. 63

1. 7621

207.61

117. 55

FlightAverage

102888

313. 24

1 7785

210. 41

118. 05

Pref l ight predic t ion

Flight performance

Fig. III-8. Stage II Engine Flight Performance

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Ill-18

Engine flight performance corrected to standard inlet conditions atthe 91FS- + 55 second time slice is shown in Table HI-10 . This iscompared with acceptance test and the predicted flight performance atstandard inlet conditions and the nominal time as used in the preflightprediction. The predicted flight performance at standard inlet condi-tions was obtained by adjusting the nominal acceptance test data for a900-poundacceptance-to-flight thrust growth obtained from analyses ofprevious Titan II and GLV flights.

TA B L EIII-10

Stage II Engine Performance Corrected to Standard InletConditions at 91FS 1 + 55 Seconds

Parameter

Thrust, chamber (Ib)

Specific impulse,engine (sec)

Mixture ratio, engine

Oxidizer flow rate,overboard (Ib/sec)

Fuel flow rate over-board (Ib/sec)

AcceptanceTest .

100, 383

312. 72

1. 8039

206. 59

114.31

Predicted Flight(including 900-Ib

thrust growth)

101,283

312. 72

1.8039

208. 54

115.34

FlightPerformance

103 ,085

312. 75

1. 8040

212. 23

117.38

d. Shutdown transient

Stage II engine shutdown was initiated by a guidance command after181. 4 seconds of burn time. The calculated shutdown impulse from91FS 0 to 91FS 0 + 20 seconds was 37, 177 Ib-sec; predicted impulse wasi &37 , 000 +7000 Ib-sec. T he impulse obtained from the + 10 g acceler-data, and illustrated by the P decay in Fig. III-9, was 25 ,658 Ib-sec,

C3using an average spacecraft/Stage II weight of 14, 325 pounds. Thiswas for the time interval from 91FS 2 to 91FS ? + 0. 631 second. Im-

pulse from 91FS 2 + 0. 631 second to 91FS 2 + 20 seconds was 11, 519 Ib-sec, utilizing the + 0. 5 gaccelerometer data and an average weight of14, 257 pounds. Tbis thrust tailoff is illustrated in Fig. Ill-10.

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1000

MDFJPS (Meas 0855)

- 1.0 +1.0 + 2 . 0Time f rom 91FS ? (sec)

Fig. III-9. S/ A 3 Shutd own Transient

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- 21

The t ime of zero thrus t occurred at approximately 91FS 9 + 26. 0

seconds. Thrust at SECO + 20 seconds was est imated at 40 pounds,wit h in the specified maximu m of 60 pounds.

e. Engine malfunction detect ion systems

The Stage II engine M D S operated satisfactori ly throughout flight.Figures III- 7 and III- 9 i l lustrate the respon se t imes and cham berp re s su re correlation during th e start an d shu td own t rans ients , respec-tively, of the malfunction detection fuel injector p re s su re switc h es(MDFJPS). The fuel in jector p re s su re i s not ins t rum ent ed and, there-fore, is not available. A sum m ary of th e significa n t switc h p a r a m e t e r sis presented in Table III- 11.

TABLE III- 11

Stage II MD FJPS Operation

P a r a m e t e r

Act ua tio n time (sec)

P at actuation (psia)

Deactuation time (sec)

P at deactu ation (psia)

91FS..+ 0.690

Invalid

91FS 2 + 0. 155

460


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- 22

. P ROP ELLANT SUBSYSTEM

1. Propellant Loading

a. Loading proced ure

A special loading in addition to the launch loading was made forG T- 7. The special loading was pe rfor m ed to corre la te propellantt ank volumes at the var ious sensor locations with previous tank cali-bration data. A flow verification test was performed pr io r to eachloading to evaluate the readiness of the FTPS for the subsequent load-ing. The op era tio n al sequen ce is given in Table III- 12.

TABLE III- 12

G T- 7 Loading Sum m ary

Opera t ion

Flow veri f icat ion

Special loading

Prelaunch flow

verificationLaunch loading

Descr ipt ion

F u e l and ox id i ze r flow up to dis-t r ibut ion area; fo rward t h roug hfou r f l owmete r s

a. Stage II oxidizer throughfour f lowmete r s

b. Dual (Stages I & I I ) oxidizerloading

c. Stage I I fue l through fourf lowmete r s

d. Stage I fue l through fourf lowmete r s

Flow up to distribution area;fo rward th roughfour f lowmete r sD u a l l oad ing

D a t e

11 November1965

16 November1965

27 November1965

3 & 4 December1965

ropeiiant loading for G T- 7 was accomplished through the tandemflowm et er system installed after the launch of G T- 5. No ser ious hard -ware problems o ccu rred du ring any of the events leading up to thelaunch; h owever , two flowm et ers were changed because out - of- toleran ceresul ts were obta ined during the special loading test and the prelaunchflow verification.

Stage I fuel m e t e r S/ N 202146 was removed from the system fol-lowin g th e special loading an d sent to M a r t in - D e n ve r for check cali-

bra t ion and acceptan ce test ing. The check calibrat ion resul ts verifiedthat the meter was in calibration an d verified resul ts obta ined du r ingthe special loading.

Flowm eter S/ N 199169, used on the Stage I fuel auxil iary positiondur ing t h e second flow ver i f ica t ion test , was remo ved because it s accu-r a c y was questionable. Check calibrat ion r e s u l t s showed a flowm et ere r r o r of 0. 1%. Flo wm eter S/ N 202146 replaced meter S/ N 199169 inthe Stage I fuel auxiliary position for the launch.

ER 13227- 7


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- 23

The t ab runs used fo r l aunch loading wer e esta blishe d by u s i n g t h edata contained in Table III- 13, and the bias derived f rom the d i f f e r encesbet wee n M a rt in - D e n ve r an d Wyle calibrat ion facil i t ies. Che cks of thetwo calibrat ion facil it ies have established th at , if a flowm ete r (fuel oroxidizer) calibrated at M ar t in - D e n v er is assumed to be c o r r e c t , then acor respon din g Wyle- ca librat ed flowm ete r wil l ind icate about 0. 3%higher. Presen t ly, it is not kn own which facility is more accu ra t e ;ho wever, the launch loadin g was based on the assumpt ion tha t theM ar tin - D enver calibrated flowm ete rs were correct. This, in effect,decreased t h e Wyle- ca l ib ra t ed f lowmet e r / t ab ru n errors r eco rdeddur ing the specia l load in gs by 0. 3% an d esta blishe d th e leas t proba -bility of payload loss.

Many combinations of possible meter errors were considered be-fore th e dec ision was m ade to bias th e load ing tab ru n s as shown inTable III- 14. The most significant cases evaluated were as follows:

Case 1: Bias all Wyle m e t e r s by 0. 3%

Wyle

Oxidizer

Wrong

Right

Right

Fuel

Wrong

Right

Wrong

Mar t in - Denver

Fuel

Right

Wrong

Right

Payload Change

0

+44

- 15

Case 2: Bias n o data (average raw results)

Wyle

Oxidizer

Right

Wrong

Right

Fuel

Right

Wrong

Right

Mar t in - Denver

Fuel

Right

Right

Wrong

Payload Change

0

- 52

- 4

In general, Case 1 is m ore de si rab le than Case 2.

A graphic display of the flowmete r- to - t ab run comparison is shownin Figs. Ill- 11 an d III- 12. In each figure, the data ar e referenced tothe tank calibrat ion made at Denver (which is synonymous to the specialloading tab run). The data for Wyle m e t e r s are not cor rec ted for thedifference between D en ver an d Wyle facilities.

CONFIDENT!ER 13227- 7


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-dtn

ao

V

2

0

N

>

QQ

""^

*

00

r

r

oCQQ

~5

5**

s3

*

3

s?+03_T*0

C

H

C

CM

Stage I oxidizer(dual loaded)+

??

5^

-"^T

r

C

H

^Stage II oxidizer

(through 4 mete

0t?3to0CO0>O

-

^Md

3

I

?

N

&

5206362

202146

3

*??

SSSQ

^^*

2 0 6 3 6 1

204278

85

??

S&

%

"

206360

2 0 2 1 6 4

Stage I oxidizer

C

+ +

S

*

199168

199167

Stage II oxidizer

.

IQ

ER1

7


79/229

III-25

TABLE III-14

GT-7 Launch Loading Tab Run Correction

Tank

Stage I fuel

Stage II fuel

Stage I oxidizerStage II oxidizer

Special LoadingAverage Error

BetweenFlowmeter and

Special LoadingTab Run (%)

-0. 11

-0. 46

+0. 24+0. 62

Bias to Accountfor Wyle Meters

in Systemfor SpecialLoading (%)

-0. 15

-0. 15

-0. 30-0. 30

Correction toOriginal Tank

CalibrationTab Run for

Launch (%)

-0. 26

-0. 61

-0. 06-0. 32

The sequence of launch propellant loading events appears in TableIII-15.

TABLE III-15

GT-7 Propellant Loading Schedule

Event

Start prechill (EST)

Start load (EST)

Hi-lite (EST)

Load complete (EST)

Stage I

Oxidizer(3 Dec1965)

2145

2210

2340

2355

Stage II

Oxidizer(3 Dec1965)

2145

2210

2248

2307

Stage I

Fuel(4 Dec1965)

0006

0030

0124

0134

Stage II

Fuel(4 Dec1965)

0006

0030

0102

0107

Mission loads for the oxidizer tanks of both stages were obtained byusing the K-factor ratio technique. This was in accord with a MartinCompany/SSD agreement that an oxidizer flowmeter tab run error ofmore than +0. 1% at hi-lite would constitute an out-of-tolerance condi-tion.

b. Total propellant loads

Total mission loads for the launch, as determined from flowmeters,are shown in Table III-16. The flowmeter totalizer readings were

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III- 2 6

Data are cor rec ted fo r ac tua l flow rates an d r e p r e s e n t the percent error of theflow meter resu l t at h i - l it e from th e original t ank calibration data.

Stage I Stage

Launch Special loading Launch Special load ing

0 161 i

206, 360

Launch | r ,tab shift '""I

" 8 199, 168 -

- 0 . 7

- 0 . 6

199, 167

- 0 . 5

- 0.4

Launch r^- 0. 3 tab shift ' \ '- 06 360

, 202, 164- 0 . 2

- 0. 1

- 0

- - 0 . 1

- - 0 . 2

0.8

- " .7

- 0 . 6

- 0.5

0.4

- 0.3

- 0.2

- 0.1

- 0

- - .i

- - 0.2

- 199, 168

- 206, 360

199, 167

- 202, 164

Note: All meters are Wylie calibrated

Fig. III- ll. GLV-7 Loading Summ ary- - Oxidizer

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- 27

Data are corrected for actual flow r a t e s an d r e p r e s e n t the percent error of theflowm ete r r e su l t a t h i - l it e f rom the or ig ina l t ank ca l ib ra t ion da ta .

Stage I Stage II

Launch Special load ing Launch Special loading

06 36W

202.146D 1 -

Launch . |- r^ v+ oK oViif t 1

- 0 . 3

- 0 . 2

- 0 . 1

- 0

j 6 3 6 1 W

- - 0. 1_ ^ - 2 04 278D

9Qg 351 VV - - 0 . 2

- - 2 0 2 . 14 6D- - 0.3

9ndn 7f l _ - Q 4 6Vi, 6 t o

- - 0.5

~ 0 . 6 - jt ab shift u x

- - 0. 7

0.3

- 0 .2

- 0. 1

- 0

- - 0 . 1

- - 0.2

- - 0.3

= 0. 4_

- . 5

- - 0 . 6

- - 0. 7

206.361W206, 362 W

204, 278 D

202, 146 D.

Legend:

W Wylie calibrat ion f lowme ter

D D enver cal ibrat ion f lowm eter

Fig. 111-12. GLV-7 Loading Summary- - Fuel

ER 13227- 7


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- 2 8

correct ed by subtract ing propellant vapo rized an d propellant rem ain -ing in th e fill l ines . Oxidizer flo wm et er loads reflec t the use of theK- fac t o r ra t io method to obtain mission loads.

Total propellant loads as de te rmined by flight verification are alsoshown in Tab le III- 16. The flight verificat ion loa ds were calculate dfrom a propellant invento ry usin g actual level sensor uncover t imesan d t ank calibrat ion data to de te rmine flow rates. Total, integrated,in - f ligh t , an d overboard propellant con sump tions were f o u n d us ing th eengine an alytical m odel. En gine start t ran sient con sum ptio n s werederived from Aero jet sum m ary r e p o r t s . Other transient propellantconsumptions an d pressurization gas weight s were calculated fromflight da t a (Tables - 37 and - 38).

The differences shown in Table III- 16 in dic at e the compar ison be -

t wee n pre flight data and postf l ight ver i f ica t ions .

TABLE - 16

Ver ificat ion of Propellant Loads

Ta n k

Stage I ox id ize r

Stage II oxidizer

Stage I f u e l

Stage I I f u e l

F l o wm e t e rI n d i c a t e d L o a d

(lb)

1 7 2 . 7 4 7 *

38,479*

9 0 , 2 0 1

2 1 , 9 8 8

R e q u e st e dL o a d(lb)

1 7 2 , 7 4 7

3 8 , 4 7 9

90, 181

21, 972

F l i g h tVe r i f i c a t i o n L o a d

(lb)

1 7 2 , 5 3 1

38, 609

90, 164

2 1 , 9 5 7

D i f f e r e n c e BetweenFl igh t Ver i f i ca t ion

a n d F l o wmeter L o a d(%)

- 0.125

+ 0 . 3 3 8

- 0.041

- 0. 141

*Miss ion load ob ta inedby K - fa c t or ratio t e c h n i q u e .

c . Propel lan t assay

Pre l aunch data from t h e propel lan t assay laboratory report (sampledon F- 4 day) for the oxidizer (n itro gen tet roxide) an d fuel (50% h ydra zin ean d 50% U D M H ) are presented in Table III- 17. Specif icat ion values arealso l i s ted . G o o d agreement i s shown between the laborato ry data andspecif icat ion r equ i r emen t s . The dens i ty was de te rmined by a pycnomete r.

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I III- 29

TAB LE - 17

Propellant Assay Summary

Fue l M IL- P - 27402 (USAF)

Hydrazine

U D M H

2

Total N H + U D M H2 4

Solids

Pa r t i c l e s on 50 m esh screen

Density (gm/cc) at 77 F

Oxidizer M IL- P - 26539 (USAF)

Nitrogen tetroxide (N 2 O 4 )

Chloride as NOC1

H 9 O equivalen t

Solids

Nonvolatile ash

Pa r t i c l e s on 50 mesh screen

Test

5 1 . 4 %

47. 8%

0.8%

99. 2%

0. 12 m g / l i t e r

0

0.9009

Test

99. 5%

#

0. 06%

0. 40 m g / l i t e r

*

*

Requ i remen t

51 + 0.9%

46. 9% min

2. 0% max

98% min

25 m g / l i t e r

0

- -

Requirement

99. 4% min

- -

0 . 2 %

10 mg/ l i t e r

- -

0

*Not repor ted .

2. Propellant Temperatures

a. Weather

A comparison of the F - 45 day prediction, the F~l day prediction

an d the actu al weat he r for the 4 December launch of G T- 7 appears inTable - 18. The F - 45 day prediction is based on wea th er for a hotDecember throu gh M arch day. The F~l day prediction was in goodagreement with the actual wea th er except for the predicted win d speed,whic h was high.

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Q

-1~1

08jfa0

1

t8a>'?Qrature6>H.aL

|

fActualfa

factualfaTfa1


85/229

- 31

b. Propellant loading temperatures

Table III- 19 com pares th e requested propellant t empera tu res at theRSV (at star t of loading) and the tank bottom probe (at hi- l ite) with theactual propel lant temperatures .

TABLE - 19

Propellant Temp erature Com parison- - RSV an d Tank Bottom Probe

System

Stage I fuel

Stage II fuel

Stage I oxidizerStage II oxidizer

RSV

Requested

26.0

26.0

20.020.0

Act ua l

26.0

26.0

20.020.0

0

0

0

0

Tank Bottom Probe

Requested

29.6

28,5

23.025.0

Act ual

28.6

29.7

23.925.5

- i .o+ 1.2

+0.9+0 .5

The requested R SV te m pe rat ur es were matched exactly, an d t ankbottom probe rea din gs were wit h in an acceptable range of accuracy.

RSV an d flowm et er tem peratu res recorded d ur in g loading are shownin Figs. - 13 and III- 14.

c . Li f to ff t e m p e r a t u r e s

A compar i son of predic ted, ac tual an d r econ s t ruc ted p rope ll an tbulk t e m p e r a t u r e s is shown in Table - 20.

TABLE - 20

Propel lant Bulk Tem peratu re Com parison

System

Stage I fuel

Stage II fuelStage I oxidizer

Stage oxidizer

F - 4 5 D ayPrediction

44.0

39.542.8

44.8

F - l D a yPrediction

48.9

4 3 . 246.4

47.1

Act ua l

4 2 . 6

4 0 . 641. 1

4 4 . 4

Recons t ruc t ed

43.8

41.041.6

43.9

Act ua l bulk t empera tu res at l i f t o f f wer e obtain ed from a compute rprogram analysis of flight data. The position of the r econs t ruc tedtem peratu res in the mixture ra t io band is shown in Figs. Ill- 15 an dIII- 16.

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4>i.

(

)a

v

a

1

ITI

I

E

1

7


87/229

-

-

II8:.

45 -.j Stage I load complete,.

r> 1,1

4 0 .

mTime o f event

3 0

25

Resume loadStage I f u e l Hi- Lite

Stage II load complete rR e su m e load

Stage I I f u e l Hi- LiteR e s u m e load

Start leak checkStart loading -

Meas 4432(Stage I If l o w m e t e r )

, ._:. Meas 4431_ (Stage I )

f l o w m e t e r )

0100

Eastern Standard Time (hr)

Fig. 331-14. Fuel Temperature During Loading

E R 13227- 7


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III- 34

6 5 P H

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: F - 4 5 days predicted: l a u n c h window

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25 j25 30 35 40 45

Bulk Fuel Tempera ture ( F)

50

Pig. HI- 15. Propellent Bulk Temperatures at Liftoff, Stage I

ER 13227- 7


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M R (pptim

M R ( m i n i m u m

"25 30 35 40 45

Bulk Fuel Tempera ture (F)

50 55

Fig. III- 16. Propellent Bulk Temperatures at Liftoff, Stage II

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- 45

3. Propellant Feed System

a. Feedl ine t rans ients

The maximum t r a n s i e n t p r e s s u r e s r ecorded at the p u m p in le t in -s t rumenta t ion bosses are listed in Table 111- 22.

TABLE III- 22

Maximum Transient P r e s s u r e s at P u m p Inlet

M e a s u r e m e n t

Stage I oxidizer (0017)

Stage I fuel (0014)

Stage II oxid izer (0510)

Stage f u e l (0507)

AtP re valveOpening

(psia)

No data

44.0

75.0

Negligible

AtIni t ial

P r e s s u r eWave (psia)

Negligible

Negligible

Negligible

Negligible

AtIgnition

(psia)

130

Negligible

*

*

AtTCV

Closing(psia)

Negligible

Negligible

65

Negligible

DesignOperat ingP r e s s u r e

(psia)

215

55

260

80

*Not available due to t e l eme t ry staging blackout.

N o dat a wer e available on the Stage I oxidizer prevalve openingpressure t r a n s i e n t . To fac i l i ta te se t t ing the ullage volum e in the oxidizers tandpipe , these valves were opened before telem etr y rec ord in g wass ta r t ed . Ignition t r ans i en t p r e s s u r e s were , in genera l, s imilar t othose on G LV- 4 an d G LV- 5 flights. Telem etry blackout norm ally ex-per ienced du rin g Stage II ignition prevents obta in ing data on sustainerengine ignition t r ans i en t s .

b. P u m p inlet suction p r e s s u r e s

Stages I and II static p r e s s u r e s at the suction m e a s u r e m e n t b o sslocat ions ar e shown in Figs. 111- 25 t h rough 111- 28, whic h presen t theprefl ight pred ic t ed , postflight r e c o n s t r u c t e d , an d best es t imate ofactual flight p r e s su re s . The postflight reco nst ru cte d curves are basedo n f l ight - m ea sure d values of ullage gas pr essure , axial load facto rs,propel lant t e m p e r a t u r e s , an d propellant loadings.

The Stage I oxidizer best estimate curve of the s ta t ic suction p r e s -s u r e s at the m e a s u r e m e n t boss (Meas 0017) co n sists of an average ofthe measured pressure an d the two oxidizer s tandpipe p r e s s u r e s(Meas 0033 an d 0034) adjusted to the M ea s 0017 boss location. TheStage I fuel suction p r e s s u r e best estimate at Meas 0014 boss lo -cation is an average of measured p r e s s u r e an d th e two fuel a c c u m u -lator p re s s u re s (Meas 0037 and 0038) adjusted to the M eas 0014boss location. The Stage II oxidizer an d fuel best est imate suctionp r e s s u r e s are the p re s s u re s m ea sure d by M eas 0510 an d 0507, re-spectively.

ER 13227- 7


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111- 46

100

90

a ,

~ 8001- -

1

70

60

50

Preflight prediction Postflight reconstruction

Best est imate o f flight suction pre s su re

87FS,

10 60 80

Time from

100 120 140 160

(sec)

Fig. HI- 25. Stage I Oxidizer Suction Pressure (Meas 001?)

ER 13227- 7


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Time from 87FS 1 (sec)

87FS,

Fig. 111 31 G T- 4 Fuel Accumula to rPiston Travel

E R 13227-7


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- 67

Dynamic f r ic t ion levels for dry a c c u m u l a t o r s were measured priorto instal lat ion of the accumulator assembl ies at Mar t in- Bal t im ore andagain pr io r to flight at Complex 19, ETR . A s u m m a r y of these frictionm e a s u r e m e n t s a p p e a r s in Table - 41 as pea k- to - pea k values ( twice theequivalent f r ic t ion fo rce in one direc t ion) .

TABLE 1 - 41Dynamic Fric t ion Levels for D ry Accu m ula t ors

S/ A

1

2

Serial No.

B011

Peak- to- Peak Fr ic t ion (psi )*

Bench

0. 9

0.6

Prefl ight

0.8

0.8

*Maximum acceptable value = 2.0 psi

Flight data do not indica te s igni f icant d i fferen ces in f r ic t ion levelsbet wee n accumula to r s .

6. POG O

The Stage I lon gitudin al oscil lat ion levels fo r th i s flight were th elowest experien ced . Flight data do no t indica te s igni f icant responsesin prop u l sion m easu remen t s un t i l imm edia t e ly p r io r t o BECO. Oxidizersuct ion p r e s s u r e (M ea s 0017) an d oxidizer standpipe p r e s s u r e s (Meas0033 an d 0034) show a buildup at the structural f requency beginning 3.5seconds be fo r e BECO. Th is buildup has been observed on previous f l ight san d is predictable analytically; i .e . , th e system gain at zero phase anglecrosses un i ty at 95% of Stage I flight t ime f rom liftoff.

Addit ion al de ta i l s on POGO appear in Chapter XII of th is repor t .

C. PRESSUR IZATION SUBSYSTEM

! . Pre launch P res su r i za t ion

At app ro xim ate ly T- 192 m inu tes, al l pro pellant tan ks were p r e s -surized, through AG E, f rom blanket p r e s s u r e level t o flight pres su relevels. The resu l t an t t im e - p r e ssu r e pro files (Fig. 111- 32) in d icat e th atth e process was n o r m a l . The t ank ullage lockup p r e s s u r e s obtainedf rom l and l ine measu remen t s m ade at T~0 an d the r e la t ed no rma l ope ra t in gp r e s s u r e r anges are presented in Table - 42.

LER 13227- 7


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- 68

:Meas 4602

' Stage I I f u e l

::

40

al- -3a

.j

id

.

M eas 4605Stage II oxidizer !

Meas 4129Stage I oxidizer

2 3

Time Afte r Initiatio n of Flight P r e s s u r e Signal (min )

Fig. 111-32. Tank Pressurization Cycle (blanket to flight pressure)

TIALE R 13227- 7


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P r e fl igh t p r e d i c t i o n Flight performance T fpnr es t imated

NOTE: AU t i m e s f rom 87FS,

0.055i

0 . 0 6 0 0 . 0 6 5 0 . 0 7 0 0 . 0 7 5I

0 . 0 8 0

Flow Ratio, W- p / Q f ( Ib / s e c p r e s s u r a n t g a s / - prope l lan t )

Pig. Ill- 37. Stage I Fuel Tank Pressurant Performance

ER 13227- 7


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270N o t e : All times f r o m 87FS 1

2 500 .12 0.13 0.14

Flow R a t io , W O p / Q o s ( I b / s e c pressurant

0 15 0. 16 0.17 0.18 0. 19

propellant)

Fig. 111-38. Stage I Oxidizer Tank Pressurant Performance

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Ill- 7 7

290

r

Flight pe rforman ce,

N o t e : All t imes f r o m 91FS50

0.10 0.11 0.12

Flow Ratio, W - p / Q - o (Ib/sec pressuran t

Fig. 111- 39. Stage II F u e l Tank Pressurant Performance

0.13 0.14 0.15 0.16 0.17ft

propellant)

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I ll - 78

TABLE 111- 44

Pressure D iffere nce Between Tank Pressure Transducer Pairs

Tank

Stage I oxidize r

Stage I fuel

Stage II oxidizer

Stage II fuel

DifferenceMaximum

(psi)

0. 15

0.35

0.40

1. 62

MeanDifference

(psi)

0.09

0.20

0.16

1.32

MaximumAllowableDifference

(psi)

1.50

1.50

2.25

2 .25

D . E N V I R O N M E N T A LC O N T R O L

1. Laun ch Veh icle Air- Con dit ion ing System

This system, which serves launch vehicle Compartment 2 and allengine start car t r idges , was operative continuously during the prelaunchactivities until vehicle liftoff. The system operated satisfactorily.Table - 45 presen t s a summary of the system param eters .

E R 13227- 7


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Ill 79

TABLE 111-45

Air-Conditioning System Performance Summary

VIeas

4403

4405

4418

4045

4046

4612

Description

GLV supplyair temper-ature

Compart-ment 2 sup-ply air massflow rate

Compart-ment 2 ex-haust airtempera-ture

Start car-tridge tem-peratureS/A 1

Start car-tridge tem-peratureS/A 2

Start car-

tridge tem-peratureS/A 3

ObservedRange

48 to50. 5 F

Approxi-mately 88Ib / min

54 to58 F

54 F (atliftoff)

53. 5 F(at lift-off)

52. 5 F

(at lift-off)

SpecifiedRange

48 to 56 F(Compart-ment 2) 48to 58 F(enginestart car-

tridges)82 Ib/min(minimum)

40 to 75 F

36 to 84 F

35 to 84 F

35 to 70 F

Remarks

Temperature of airsupplied to GLVCompartment 2 andthe engine startcartridges

.

Manual hold param-eter

S/N 0636763Manual hold param-eter

S/N 0859574Manual hold param-eter

S/N 0859190

Manual hold param-eter

ER 13227-7


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IV- 1

IV FLIGHT CONTROL SYSTEM

An alysis of the G T- 7 Flight Control System (FCS) m easure d pa ram -eters indicated satisfactory system ope ration du ring both Stages I and flights. The pr imary FCS was in command throughout and no swit ch -over to the secon dary system was required.

A. STAG E I FLIGHT

1. Ignition and Liftoff Transients

Peak actuator travels an d rate gyro d isturban ces record ed duringthe ignition an d ho lddown per iod are presented in Table I V- 1.

TABLE IV- 1

Transients During Stage I Holddown Period

Act ua to rDesignation

Pitch, lj

Yaw / rol l , 2j

Yaw /roll , 3j

Pitch, 4 X

Axis

Pitch

Yaw

Roll

Maximum During IgnitionTravel

(in.)

- 0.094

- 0.090

+ 0.200

- 0.071

Time from LO(sec)

- 2 .75

- 2 .75

- 2 .80

- 2 .80

Maximum Rafe, Stage IG yro (deg/ sec)

Pr imary

- 0. 20

+ 0.39

+ 0.38

Secondary

- 0.30

+ 0.19

+ 0.40

Maximum DuringHolddown Null Check

( in . )

+0.02

- 0.04

- 0.02

- 0. 01

Maximum Rate , Stage Gyro (deg/sec)

Pr imary

+ 0.39

- 0.19

- -

Secondary

+0.47

+ 0.19

- -

The combination of thrust misalignment and en gine misalignment

at full thrust initiated a roll t ransient at liftoff. The response of theFCS to correct the offset kept the roll rate to a maxim um of 0. 9 dee /sec cou nt er- clockwise ( C C W ) at 0. 22 second after liftoff (Fig. IV- 1).The rate oscillation had a basic frequency of 5. 4 r a d / s e c , damping outin 1.7 seconds. As shown on the roll error curve in Fig. IV- 1, a rollbias of 0. 16 degree CCW was in tr od uced a t liftoff by the equivalent en -gine misalignment of 0. 04 degree.

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IV- 2

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U - 1. 0i 0. 5 1.0 I. : 2. 0 2. 5 3. 0 3

Time f rom Liftoff (sec)

Fig. IV 1. Lifto ff Boll Transients

E R 13227- 7


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iv-

2. Roll an d Pitch P r og ram s

The TARS roll an d pitch programs performed nominally as shownin Table IV- 2. The 1. 04- second roll program was of insufficient dura-tion for the vehicle to reach steady sta te ; ho wever, the torquer m onitorindicated that the proper program was achieved. The maximum rollan d pitch overshoots which occur red at the initiation of their respectiveprograms were 1. 6 d e g / s e c CW for roll an d 0. 9 d e g / s e c down for pitch.

TABLE IV- 2

TARS Roll an d Pitch Prog rams

Program

Roll

StartStop

Pitch Step 1Start

Pitch Step 2Start

Pitch Step 3StartStop

Time fromLiftoff(sec)

19.3520.39

22.94

88.04

118.66162.02

NominalTime(sec)

19.4420.48

23.04

88.32

119.04162. 56

Rate GyroAverage(deg/ sec)

- 0. 71

- 0. 50

- 0.23

Torquer MonitorIndication(deg/sec)

+ 1.26

- 0.69

- 0.50

- 0.25

Nominal

Rate(deg/sec)

+ 1.25

- 0.709

- 0.516

- 0.235

3. TARS- IGS Comparison (Stage I)

The TARS and IG S att i tude error signals during Stage I flight forthe pitch, yaw an d roll axes are presented in Figs. IV- 2, IV- 3 an d IV- 4.The dispersion bet ween th e TARS an d IG S attit ud e was caused by a c o m -bination of TAR S gyro and IGS- IMU dri f ts , errors in open loop guidan ceprograms an d r e f e r e n c e axis c r o ss- c o u p l i n g . The dispers ion (TARSatt itud e m in u s IG S att itud e) at BECO was - 0. 78 degree in the pitch axis,+ 0. 31 degree in the yaw axis an d + 0. 10 degree in the rol l axis.

4. Stage I Flight Disturbances

Veh icle disturban ces du rin g Stage I flight wer e ca used by the pre-vailing win ds a loft. The flight control system response to these dis-tu rban ces was n orm al and well con tro lled. The yaw com ponen t of wind

ER 13227- 7


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IV- 4

+ 2 . 0

+ 1.0

S ?

1

. I I

+ 2 . 0

+ 1.0-

60 80 100Time from Liftoff (sec) BECO

Fig. IV-2. Pitch Attitude Error History During Stage I Flight

ER 13227- 7


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IV-5

+ 2.0

-

:

-:

-I

B

9

+2.0

20 60 80 100Time from Liftoff (sec)

120 140 J160

BEC O

Fig. IV-3. Yaw Attitude Error History During Stage I Flight


139/229

I V- 6

.a

- .-

- . ':t

20 40 60 80 100

Time from Liftoff (sec)

120 140 160

BECO

Fig. IV- U. Ro ll Att itu de Error History During Stage I Flight

ER 13227- 7


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IV- 8

derived from the Ra winson de Run A N / G M D - T - 1Hour Cape Kenn edywin d profile compares very favorab ly with th e yaw att itude er rorcurves in Fig. IV 3. D uring these win d disturbances, oscillationsbetwee n 1. 0 and 1. 6 ra d / sec with a p eak- to - peak overshoot amplitudeof less than 0. 3 degree of attitude er ror occ urr ed in pitch an d yaw atthe predicted GT- 7 rigid body oscillatory mode frequen cies, whichvaried with flight con dition . The maximum rates an d attitude errorsrecor ded du ring Stage I flight are shown in Table IV 3. Since th e levelof pitch and yaw excitation was of high ma gnitude, causing up to 2 de-grees peak- to- peak pitch and yaw attitude er rors during th e max qregion, th ere was inertial coupling which pro du ced excitat ion on theroll channel.

The time for FCS gain change on G T- 7 was changed to LO + 110

secon ds on the basis that inadequate stability margins would exist forth e previous gain change time (LO + 105 seconds). At th e actual timeof gain chan ge (LO + 109. 6 seconds), th ere was a pitch attitude er rorof 0. 45 degree no se up . The small but highly damped pitch tran sien treached a maximum of 0. 75 degree nose up. Preflight stability calcu-lations indicat e that , with an e r ro r of 0. 45 degree, the pitch transientfor gain change at LO + 110 second s would overshoot to a maximum of0. 79 de gree pit ch er ror. The reduction in gains reduc ed t he amoun t ofengine deflection, thus causing the transient to occur.

Since the er ror in yaw at th e tim e of gain chan ge was almost zero,there was no noticeable resultant transient.

Ana lyses indicate th at the cont rol system react ed pro perly to theflight conditions in existen ce both before an d after gain change.

5. Stage I Static G ains

The prim ary FCS static gains as determ ined from telemet ry datawere within the instrumentation inaccuracy of preflight evaluations an dindicate that no static gain deterioration was experienced during Stage Iboost flight.

. STAG E II FLIGHT

1. Staging Transients

During staging, m od erat e sustain er vehicle ra tes an d att i tude e r r o r swer e ob served . The maximum att i tude e r ro r s , measured from the pre-BECO level, and the m aximu m vehicle ra tes are given in Table I V ~ 4 .

ER 13227- 7


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IV-9

TABLE IV-4

Maximum Staging Rates and Attitude Errors

Axis

Pitch

Yaw

Roll

Rates (deg /sec)Primary

+2.39

-2.43

+1.77

-0.69

+1. 18

-2.00

Axis

Pitch

Yaw

Roll

Secondary

+2.82

-2.32

+1.70

-0.90

+1.22

-2.24

Attitude Error(deg)

-1.37

+1.78

-1. 10

Flight Time (sec)Primary

155.77

156.32

156.88

159.26

157. 14

155.98

Secondary

155.77

156.32

156.96

156.01

157.01

156.38

Flight Time(sec)

157.9

158.0

156.7

2. Stage II Attitude Errors and Biases

Pitch and yaw attitude errors are shown in Figs. IV-5 and IV-6;after the staging transient, the roll attitude error remained constantat -0. 14 degree. The predicted pitch and yaw attitudes are for thecenter-of-gravity displacement from the vehicle longitudinal axis andthe position of the roll thrust off the longitudinal axis. The additionalbiases from the predicted attitudes, -0. 85 degree in pitch and +1. 25degrees in yaw, are caused by engine thrust vector misalignment dueto structural deformation at the engine gimbal assembly. These biasesare of the same magnitude noted on previous flights and are within pre-dicted limits. The deviation from the biased predicted attitudes is due

to system hysteresis and gain sensitivity.

3. Response to Radio Guidance Commands

The TARS tinier generated the guidance enable command at LO +161. 99 seconds. Response to the first pitch command was at LO +168. 37 seconds and consisted of a small down command followed bya full 2.0 deg/sec pitch-down command for 2.0 seconds. The remain-der of the pitch commands was less than 0. 25 deg/sec. Response to

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XII- 1

XII. AIR FR AM E SYSTEM

A. STRUCTURAL LOADS

An alysis of G T- 7 flight data in dicat es that the loads e xperien cedwere well within the s t ruct ura l capabi li t ies of th e launch vehicle. Themost cr i t i ca l loading occurred, characteristically, at - BECO wherethe load aft of Station 320 reached 98. 5% of design limit load in com-press ion . Instrumentation for dynamic response data consisted of rategyros fo r lateral dynamic loads an d axially mounted acce le romete r sfor longitudinal dynam ic loads. No major anomalies affecting the air-fram e occ urred du ring flight; unu sually high am plitude propellant sloshoscil lations occurred during Stage II midflight bu t were n ot consideredto be detrimental .

1. Preignit ion

The 1 g dead weight distribut ion is the only con tribu tio n to steadyaxial loading in the preignition period. G round winds were approximately14 m ph from a direction of 320 degrees (cri t ical win d azimuth for groundwinds), whic h prod uced steady bending of 220, 000 in. - Ib and win d inducedoscil latory (WIO) oads of 765, 000 in. - Ib (Fig. XII- 1) at Station 1224.The WIO response represents approximately 40% of th e WIO design l imitbending m om en t; Table XII- 1 shows the compar ison of a ll G LV WIOexperience to date.

TABLE XII- 1

Comparison of G LV WIO Loads

Flight

G T- 1

GT- 2

GT- 3

GT- 4

GT- 5G T- 7

WIO Load at Station 1224(% of WIO design l imit bending m o m e n t )

52

5

29

3

2

40

2. Launch Prerelease

Ignition t ran sients were norm al, and the attendant dynam ic axialloads as measured by the BLH system are shown in Fig. XII- 2 togetherwith th e stead y axial load. The prerelease lateral dynamic loading wasdue to the combined effects of ground winds and, most significantly,engine start t r ans ien t s ; this loading is shown in Fig. XII- 3.

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XII-2

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Fig. XII-1. Bending Moments Due to Ground Winds: Preignition

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XII- 3

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XII- 4

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XII- 5

3. Launch Pos t re lease

A com parison (Table XII- 2) of the G T- 7 liftoff load factor with t hoseof previous laun ches ind icat es that this flight exper ienced the lowestin i t ia l steady accelera t ion to date. This can be at t r ibuted to th e lowthrus t class Stage I engine coupled with increased propellant an d space-craf t weigh ts.

TABLE XII- 2

Comparison of G LV Liftoff Load Factors

Flight

G T- 1G T- 2

G T- 3

G T- 4

G T- 5

GT- 7

Liftoff Load Factor(g)

1.27

1.27

1.27

1.27

1.28

1.26

Dynamic defo rmat ion m odes in evidence at pos t re lease consisted ofthe first an d f i f th s t ruc tu ra l bending and Stage I engine mod es in the

l a t e r a l plane and the f irst axial mode in the longitudin al d i rec t ion.Frequency correlat ion between calculated an d observed modes during thethe flight is given in Fig. XII- 4; th e result ing dynamic bending m o m e n tin the postre lease condit ion is shown in Fig. XII- 5.

4. Stage I Flight

The most signif icant periods of Stage I flight for ai r f rame loadingoccu rred at M ax C N q a an d at pre- BECO. On th is flight , M ax C N qa

a aoccurred at LO + 80 seconds, slightly later in flight than on previousvehicles. A 25- fps win d shear spike at an al t i tude of 44, 000 feet ac-

counted fo r this late occur rence .In com paring loads at LO + 80 secon ds with loads at the t r ad i t iona l

LO+ 69 second M ax qa flight time, several interesting observations. ca n be made:

(1) All of th e lateral dynamic responses noted at LO + 69 secondswere also observed at LO - I - 80 seconds, and,

Launch Vehicle No. 7 Flight Evaluation

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Transcript of Launch Vehicle No. 7 Flight Evaluation